Apparatus for power generation with low drag rotor and ramjet assembly

ABSTRACT

An apparatus ( 100 ) for generation of mechanical and electrical power. Ramjet type thrust modules ( 102   a   , 102   b ) operate at supersonic speeds (preferably Mach 3 to 4) at the distal or tip ends ( 116   a   , 116   b ) of a low aerodynamic drag rotor ( 106 ). Rotor ( 106 ) is affixed at a hub means ( 114 ) to a power output means including central rotating upper ( 104   a ) and lower ( 104   b ) shaft portions. Rotor ( 106 ) is a structural member which transmits the thrust generated by the thrust modules ( 102   a   , 102   b ) to the shaft portions ( 104   a   , 104   b ). The ramjet thrust modules ( 102   a   , 102   b ) capture and compress a supplied free air stream, which is mixed with and oxidizes a convenient liquid or gaseous fuel such as natural gas from fuel supply means ( 103 ). Combustion gases expand to create thrust to rotate the thrust modules ( 102   a   , 102   b ), which are constrained by the rotor ( 106 ), to rotates about the axis defined by the shaft ( 104   a   , 104   b ) at supersonic thrust module velocities, producing shaft energy. Escaping exhaust gases ( 160 ) may be cooled by passing them through an enthalpy extraction section ( 162 ) to heat a secondary heat transfer fluid ( 166 ). If the secondary heat transfer fluid ( 166 ) is water, the steam may be used directly for its thermal energy, or the steam sent to a steam turbine to produce additional shaft energy. Combustion gases ( 160 ) may also be directed through a reaction turbine ( 1002 ) to utilize remaining kinetic energy to generate shaft energy. The apparatus and method is particularly useful for generation of electrical and mechanical power at substantially improved efficiency rates when compared to conventional, prior art power plants.

This is a divisional of application(s) Ser. No. 07/945,228 filed on Sep.14, 1992 now U.S. Pat. No. 5,372,005, and a C-I-P of InternationalApplication PCT/US93/08713 filed on Sep. 14, 1993 and which designatedthe U.S.

A portion of the disclosure of this patent document contains materialwhich is subject to copyright protection. The owner has no objection tothe facsimile reproduction by anyone of the patent document or thepatent disclosure, as it appears in the Patent and Trademark Officepatent file or records, but otherwise to reserves all copyright rightswhatsoever.

TECHNICAL FIELD OF THE INVENTION

My invention relates to a novel, revolutionary apparatus and method forthe generation of electrical and Mechanical power. More particularly, myinvention relates to a power plant driven by thrust modules, which arepreferably ramjet engines, and to novel rotors designed to withstand theextremely high tensile stress encountered while rotatably securing suchthrust modules. The rotors are design ed for operation at supersonic tipspeeds while maintaining low aerodynamic drag, and are constructed ofcomposite carbon fiber and/or metal matrix composites. Power plants ofthat character are particularly useful for generation of electrical andmechanical power at substantially improved efficiency rates whencompared to various conventional power plant types.

BACKGROUND OF THE INVENTION

A continuing demand exists for a simple, high efficiency, inexpensivepower plant which can reliably provide electrical and mechanical power.A variety of medium size electrical or mechanical power plants couldsubstantially benefit from a prime mover which provides a markedimprovement in overall efficiency. Such medium size mechanical orelectrical power plants—in the 10 so 100 megawatt range—are required ina wide range of industrial applications, including rail locomotives,marine power systems, aircraft engines, and stationary electric powergenerating units. Power plants in this general size range are also wellsuited to use in industrial cogeneration facilities. Such facilities areincreasingly employed to service industrial thermal power needs whilesimultaneously generating electrical power.

Power plant designs which are now commonly found in co-generationapplications include (a) gas turbines, driven by the combustion ofnatural gas, fuel oil, or other fuels, and capturing the thermal andkinetic energy from the combustion gases, (b) steam turbines, driven bythe steam which is generated in boilers from the combustion of coal,fuel oil, natural gas, solid waste, or other fuels, and (c) large scalereciprocating engines, usually diesel cycle and typically fired withfuel oils.

Each of the aforementioned types of power plants are complex integratedsystems. Such plants often include many subsystems and a large number ofindividual parts. The parts often must be manufactured to exactingdimensional and mechanical specifications. As a result, such powerplants are relatively expensive to manufacture, to install, and tooperate. Also, in the event of failure of a part or subsystem, therequired repairs are often quite expensive. Frequently, repairs mayrequire substantial disassembly of subsystems to gain access toindividual parts, in order to repair or replace the faulty componentsand return the plant to an operational condition.

Of the currently available power plant technologies, diesel fueledreciprocating and advanced turbine engines have the highest efficiencylevels. Base efficiencies are often in the range of 25% to 40%, based onnet work produced when compared to the energy value of the fuel source.Unfortunately, at power output levels greater than approximately 1megawatt, the size of the pistons and other engine components requiredby reciprocating engine systems become almost unmanageably large, and asa result, widespread commercial use of larger sized reciprocating enginesystems has not been accomplished.

Gas turbines perform more reliably than reciprocating engines, and aretherefore frequently employed in plants which have higher power outputlevels. However, because gas turbines are only moderately efficient inconverting fuel to electrical energy, gas turbine powered plants aremost effectively employed in co-generation systems where, as mentionedabove, both electrical and thermal energy can be utilized. In that way,the moderate efficiency of a gas turbine can in part be counterbalancedby increasing the overall cycle efficiency.

Fossil fueled steam turbine electrical power generation systems are alsoof fairly low efficiency, often in the range of 30% to 40%. Such systemsare commonly employed in both utility and industrial applications forbase load electrical power generation. This is primarily due to the highreliability of such systems. However, like gas turbine equipment, steamturbine equipment is most advantageously employed in situations whereboth mechanical and thermal energy may be utilized, thus increasingoverall cycle efficiency.

Because of their moderate efficiency in conversion of fuel input toelectrical output, the most widely used types of power plants, namelygas turbines and combustion powered steam turbine systems, depend uponco-generation in industrial settings to achieve advantageous commercialelectricity cost levels. Thus, it can be appreciated that it would bedesirable to be able to generate electrical power at higher overallefficiency rates than is commonly achieved today, especially whencompared to the currently utilized gas and steam turbine based powerplants.

THE PRIOR ART

Ramjets are widely know and have been utilized, primarily in aerospaceapplications, since the 1940s. Basically, a ramjet is a fixed geometrycombustion chamber which is propelled through an airstream by the thrustreaction of the chamber against escaping combustion gases which havebeen generated by oxidizing an injected fuel with the incoming airsupply. The configuration of ramjet engine inlets, fuel injectionrequirements, combustion chamber configurations, and ignitionrequirements have been the subject of much study and technicaldevelopment over many years.

Early ramjets were described, for example, in German Patent No. 554,906, issued Nov. 2, 1932 to Ing. Albert Fono. Ramjets have also beenexperimentally employed to assist in the rotation of helicopter bladesabout a central shaft. For example, see the National Advisory Committeefor Aeronautics (NACA) research memorandum (NACA RM L53DOZ) for a ramjetpowered helicopter rotor. However, insofar as I am aware, ramjets havenot been employed in commercial power plants for production ofelectricity.

SUMMARY OF THE INVENTION

I have now invented, and disclose herein, a novel, revolutionary powergeneration plant design. My power plant design is based on the use of aramjet engine as the prime mover, and has greatly increased efficiencieswhen compared to those heretofore used power plants of which I am aware.Unlike most power plants commonly in use today, my power plant design issimple, compact, relatively inexpensive, easy to install and to service,and otherwise superior to currently operating plants of which I amaware.

My novel power plants have a unique low aerodynamic drag rotor portion.The rotor is constructed of metal matrix composites and/or high strengthcarbon fiber, and can be operated at rotating speeds well above thosewhich would induce tensile and compressive strains that would causeconventional materials such as steel or titanium to fail.

Thus, the rotor design used in my power plant overcomes two importantand serious problems: First, at the supersonic tip speeds at which mydevice operates, the rotor design minimizes aerodynamic drag, thus itminimizes parasitic losses to the power plant due to the drag resultingfrom the movement of the rotor in an airstream. Second, the compositedesign provides the necessary tensile and compressive strength, whereneeded in the rotor, to prevent internal separation of the rotor byvirtue of the centrifugal and centripetal forces acting on the rotormaterials.

Solving the two aforementioned problems are critical elements of myinvention. Operation of a rotary ramjet driven rotating power generationapparatus at the supersonic tip speeds considered desirable forefficient operation would be impossible with conventional constructionmaterials such as high strength steel. Also, it is important that apower plant avoid large parasitic losses that undesirably consume fueland reduce overall efficiency.

I have now developed novel rotor designs for use in combination with aramjet driven power generation system, so as to enable high speed,aerodynamically efficient rotor operation. In one embodiment, a biplanerotor includes an upper triangularly shaped portion and an opposinglower triangularly shaped portion; the upper and lower portions both aresecured to and extend from opposing sides of a central hub portion. Thecentral hub portion is rotatably secured in an operating position alongan axis formed by upper and lower shaft portions. The upper and lowerrotor portions are situated so that air may pass above the upper portionand below the lower portion. More importantly, air may pass through agap between the upper and lower portions with minimal aerodynamic drag.

Attached to the distal end of each pair of rotors are ramjet enginethrust modules. The inward edge portion of the ramjet engine attaches tothe outer edge portions of the opposing upper and lower rotor portions,thus affixing the ramjet engine to the rotor. In various embodiments,the ramjet may be further secured to the rotor by either an externalendcap or by externally wound composite fiber bundles.

The ramjet engines are situated so as to engage and to compress thatportion of the airstream which is impinged by the ramjet upon itsrotation about the aforementioned shaft. I have also provided in mydesign a feature to insure that a relatively clean airstream (free ofthe rotor's own wake turbulence) will be encountered by the rotatingrotor and ramjet. This is accomplished by circulating, generally alongthe aforementioned axis of rotation, an airstream which can both replacethe gases scooped up by the ramjet compression as well as sweep away thewake from the just turned rotor.

Fuel is added to the air which has been compressed in the ramjet inlet.The fuel may be conveniently provided to the ramjet engine combustionchamber through use of fuel supply passageways communicating between theramjet and a fuel source. Fuel passageways may allow fuel flow upwardlyfrom the bottom shaft portion and downwardly from the top shaft portion,then such passageways are turned outwardly through the hub portion andthence radially outwardly through either or both of the rotor portions,then on through the outer edge portions, and thence through fuelinjection ports to the ramjet engine combustion chambers. The combustiongases formed by oxidation of the fuel escape rearwardly from the ramjet,thrusting the ramjet tangentially about the axis formed by the shaftportions, thus turning the rotor and the shaft portions. The power sogenerated by the turning shaft may be used directly in mechanical form,or may be used to power an electrical generator and thus generateelectricity.

In one embodiment, the outlet portions of the ramjet are positioned sothat the combustion gases may impinge on a set of heat transferelements, so as to cool the combustion gases by way of heating up a heattransfer fluid such as water which is circulated within the heattransfer elements. Ultimately, the cooled combustion gases may beexhausted to the ambient air.

In another important embodiment, an annular reaction turbine isadditionally provided surrounding the exit to the exhaust gas heatexchanger. This annular reaction turbine captures the substantialkinetic energy remaining in the exhaust gas flow, so as to improveoverall cycle efficiency.

In yet other embodiments, the rotor may be provided in the shape of adisk, discus, or similar shape.

In yet another embodiment, a small central disk rotor with outwardlyextending upper and lower biplane rotor portions are provided.

Other embodiments provide further variations in the air flowconfiguration and in provision of the fuel supply means.

In addition to the foregoing, my novel devices are simple, durable, andrelatively inexpensive to manufacture.

OBJECTS, ADVANTAGES, AND FEATURES OF THE INVENTION

From the foregoing, it will be apparent to the reader that one importantand primary object of the present invention resides in the provision ofnovel, improved mechanical devices to generate mechanical and electricalpower.

More specifically, an important object of my invention is to provide aramjet driven power generation plant which is capable of withstandingthe stress and strain of high speed rotation, so as to reliably providea method of power generation at a very high efficiency rate.

Other important but more specific objects of the invention reside in theprovision of power generation plants as described in the precedingparagraph which:

allow the generation of power to be done in a simple, direct manner;

have a minimum of mechanical parts;

avoid complex subsystems;

require less physical space than existing technology power plants;

are easy to construct, to start, and to service;

have high efficiency rates; that is, to provide high heat and high workoutputs based on heating value of fuel input to the power plant;

in conjunction with the preceding object, provide lower power costs tothe power plant operator and thus to the power purchaser than ispresently the case;

cleanly burns fossil fuels;

in conjunction with the just mentioned object, results in fewer negativeenvironmental impacts than most power generation facilities currently inuse;

have a fuel supply design which efficiently supplies a ramjet;

have a rotating element with a structure able to withstand the stressesand strains of rotating at very high tip speeds; and which

have a rotating element design which provides operation with minimalaerodynamic drag.

A feature of one embodiment of the present invention is the use of anovel biplane rotor which provides minimal aerodynamic drag at the highrotational design speeds, thereby enabling the power plant to minimizeparasitic losses, with the resulting advantage of high overall cycleefficiencies.

Another feature of the present invention is the use of a monofilamentcarbon fiber winding as an integral part of the structure of the rotor,which provides the advantage of high strength, thus enabling operationat rotational speeds above stress failure limits of conventionalmaterials such as steel and titanium.

Other important objects, features, and additional advantages of myinvention will become apparent to those skilled in the art from theforegoing and from the detailed description which follows and theappended claims, in conjunction with the accompanying drawing.

BRIEF DESCRIPTION OF THE DRAWING

In the drawing, identical structures shown in the several figures willbe referred to by identical reference numerals without further mention.Also, closely related structures in the several figures may be given thesame number but different alphabetic suffixes.

FIG. 1 shows a cross section of the power plant apparatus, including theramjet thrust module, a biplane rotor, a central hub, a rotating shaft,and air flow ducts. Additionally, fuel supply lines, exhaust heatrecovery equipment, a primary generator, the starter motor and a gearboxare illustrated.

FIG. 2 is a horizontal cross section of the power plant apparatus, takenthrough line 2—2 of FIG. 1, more clearly showing the location of theramjet thrust modules, a two-armed biplane rotor, the central hub andshaft, air flow ducts, and the heat transfer elements used for coolingexhaust gases.

FIG. 3 is an enlarged detail of the power plant apparatus similar tothat first shown in FIG. 1, showing in enlarged detail the ramjet thrustmodule, a biplane rotor, a central hub, a rotating shaft, air flow ductsystem, and the heat transfer elements for cooling exhaust gases.

FIG. 4 is an enlarged detail of a power plant similar to the oneillustrated in FIG. 3; however, in FIG. 4, the plant does not include aheat transfer section for cooling exhaust gases.

FIG. 5 is a partial isometric view of a biplane rotor of the typeprovided in the power plant apparatus of FIG. 1 above. This figure showsthe central hub structure, the upper and lower biplanes, the carbonfiber windings located inside the biplanes, and the ramjet thrust modulemounted at the distal end of the biplanes.

FIG. 6 is a vertical cross-sectional view taken through line 6—6 of FIG.5, showing the cross-sectional structure of the biplane rotor, as wellas the location of the ramjet thrust module at the distal end of therotor.

FIG. 7 is a perspective view of the distal end of a biplane rotor,showing a ramjet thrust module attached thereto.

FIG. 8 is a horizontal cross section, taken through line 8—8 of theramjet thrust module of FIG. 7, looking downward at the construction ofthe thrust module.

FIG. 9 is a vertical view, looking rearward in the direction of theexhaust in the thrust module of FIG. 7, taken at the station indicatedby line 9—9 in FIG. 8. This view shows the interior air flow path of thethrust module.

FIG. 10 is a vertical cross-sectional view, looking rearward in thedirection of the exhaust, cut through the thrust module of FIG. 7, takenat the station indicated by line 10—10 of FIG. 8. This view shows thethickening wall portions of the thrust module at this station, as wellas the air flow path already seen in FIG. 9 above.

FIG. 11 is a vertical cross-sectional view, looking rearward in thedirection of the exhaust, cut through the thrust module of FIG. 7, takenat the station indicated by line 11—11 of FIG. 8. This view shows theouter cap of the thrust module, as well as the first layer of thereinforcing carbon fiber windings which wrap around the end of thethrust module.

FIGS. 12 through 15 are vertical cross-sectional views, looking rearwardin the direction of the exhaust, cut through the thrust module of FIG.7, taken at the stations indicated by reference of FIG. 8, similar toFIGS. 9 through 11 above. FIGS. 12 through 15 show the varying thicknessof the reinforcing carbon fiber windings, as well as the shape of theinterior of the thrust module air flow path.

FIG. 16 is a vertical cross-sectional view, looking rearward in thedirection of the exhaust, cut through the thrust module of FIG. 7, takenat the station indicated by line 16—16 of FIG. 8. This view shows theouter cap of the thrust module, as well as the shape of the interior ofthe thrust module air flow path, at this point in the exhaust section.

FIG. 17 is a vertical view, looking forward in the direction of the airinlet from the rear of the thrust module of FIG. 7, taken at the stationindicated by line 17—17 of FIG. 8.

FIG. 18 is a horizontal cross-sectional view, similar to the view firstset forth in FIG. 8 above, showing a first alternate configuration forthe interior of a ramjet thrust module, utilizing a reverse Lavalinternal contraction nozzle. The figure also shows areas requiringcarbon fiber reinforcement for operation in the present invention.

FIG. 19 is a horizontal cross-sectional view, similar to the view firstset forth in FIG. 8 above, showing a second alternate configuration forthe interior of a ramjet thrust module, utilizing a mixed contractioninlet nozzle. The figure also shows areas requiring carbon fiberreinforcement for operation in the present invention.

FIG. 20 is a horizontal cross-sectional view, similar to the view firstset forth in FIG. 8 above, slowing a third alternate configuration forthe interior of a ramjet thrust module, utilizing an ejector augmentedflow path. The figure also shows areas requiring carbon fiberreinforcement for operation in the present invention.

FIGS. 21A, 21B, and 21C illustrate air flow spillage and shock wavelocation for the startup of a mixed contraction inlet ramjet thrustmodule. The mixed contraction inlet is similar to the second alternatethrust module configuration first illustrated in FIG. 19 above. Thefigure also shows areas requiring carbon fiber reinforcement foroperation in the present invention.

FIG. 21A shows shock wave location and spillage for operation of aramjet thrust module well below design mach number. The figure alsoshows areas requiring carbon fiber reinforcement for operation in thepresent invention.

FIG. 21B shows shock wave location and spillage for operation of a mixedcontraction inlet slightly below design mach number. The figure alsoshows areas requiring carbon fiber reinforcement for operation in thepresent invention.

FIG. 21C shows the shock wave location and the captured airstream tubeas would be present in the operation of a mixed contraction ramjetengine at design mach number. The figure also shows areas requiringcarbon fiber reinforcement for operation in the present invention.

FIG. 22 illustrates the airflow configuration for an internalcontraction inlet ramjet. The figure also shows areas requiring carbonfiber reinforcement for operation in the present invention.

FIG. 22A illustrates a generalized cross section configuration of aninternal contraction type ramjet thrust module, similar to that firstillustrated in FIG. 8 above. The figure also shows areas requiringcarbon fiber reinforcement for operation in the present invention.

FIG. 23 illustrates the airflow configuration for a self-starting, mixedcompression inlet ramjet thrust module.

FIG. 23A illustrates a generalized cross section configuration of amixed compression type ramjet thrust module, similar to that firstillustrated in FIG. 9 above.

FIG. 24 shows a generalized cross section configuration of an internalcompression type ramjet thrust module, similar to that first illustratedin FIG. 8 above, now showing the combustor location in the thrustmodule, as well as describing other regions of the nozzle.

FIG. 25 shows a generalized cross section configuration of an internalcompression type ramjet thrust module, similar to that first illustratedin FIG. 24 above, further describing the various regions of the thrustmodule.

FIG. 26 shows in graphical form the variation in thrust output from thethrust module at various throttle settings, for a design Mach number of3.5.

FIG. 27 shows in graphical form the variation in thrust module thrust atvarious Mach numbers.

FIG. 28 is an illustration of the shock structure and surface pressuredistribution resulting from the mechanical deflection of a supersonicflowfieid by a diamond shaped cross section, such as a diamond shapedrotor section.

FIG. 29 is an illustration of the shock structure and surface pressuredistribution resulting from the mechanical deflection of a supersonicflowfield by a bi-convex shaped cross section, such as a bi-convexshaped rotor section.

FIG. 30 is an illustration of the shock structure and surface pressuredistribution resulting from the mechanical deflection of a supersonicflowfield by a lifting flat plate, such as a flat plate rotor section.

FIG. 31 is an illustration of the attenuation of shock waves from adiamond shaped cross section, through interaction with expansion waves.

FIG. 32 is an illustration of the attenuation of shock waves from a flatplant shaped cross section, through interaction with expansion waves.

FIG. 33 illustrates pressure drag reduction due to shock cancellationwithin a biplane rotor.

FIG. 34 illustrates shock cancellation within a biplane type rotor whenthe biplane is not operating at the design mach number.

FIG. 35 illustrates a desirable rotor geometry, and shows variance ofbiplane height and internal gap height in the biplane rotors to achievelow aerodynamic drag.

FIG. 36 shows the flowfield near a flat disc rotating in a quiescentfluid.

FIG. 37 shows the variation in the moment required to spin a flat discin a quiescent fluid. The moment required is expressed in terms of adimensionless moment coefficient. The figure shows the theoreticallypredicted behavior for laminar and turbulent flows at various rotationalReynolds numbers.

FIG. 38 is a schematic representation of a disc rotating in a housing.

FIG. 39 is a schematic representation of a disc rotating in a housing,taken along line 39—39 of FIG. 38.

FIG. 40 shows the variation in the moment required to spin a Flat discinside a housing. The moment required is expressed in terms of adimensionless moment coefficient. The figure shows the theoreticallypredicted behavior for laminar and turbulent flows at various rotationalReynolds numbers as well as a comparison to the moment coefficientsrequired to turn a disc without a housing set forth in FIG. 37 above.

FIG. 41 is a perspective view of a first alternate embodiment of therotor of the present invention, here shown as a flat disk.

FIG. 42 is a vertical cross-sectional view of the flat disk rotor firstillustrated in FIG. 41.

FIG. 43 is a perspective view of a second alternate embodiment of therotor of the present invention, here shown as a tapered disk.

FIG. 44 is a vertical cross-sectional view of the tapered disk rotorfirst illustrated in FIG. 43.

FIG. 45 is a perspective view of a third alternate embodiment of therotor of the present invention, here shown as a small central disk withan outwardly extending biplane portion.

FIG. 46 is a vertical cross-sectional view of the combinationdisk/biplane rotor first illustrated in FIG. 45.

FIG. 47 illustrates, for purposes of stress analysis and comparison, aslender rod rotating about an axis perpendicular to its own longitudinalaxis.

FIG. 48 illustrates, for purposes of stress analysis and comparison, aflat disc of uniform thickness rotating about an axis perpendicular toits own plane.

FIG. 49 is a graph showing the variation of specific stress withrotation rate, for both a non-tapered slender rod and for a rotatingdisc.

FIG. 50 is a graph which illustrates a desirable rotor taper schedulefor stress reduction, i.e., the variation of the rotor cross-sectionalarea versus radial position.

FIG. 51 shows the spanwise variation in radial stress in the biplanegutters.

FIG. 52 shows the spanwise variation in radial stress in carbon filamentwindings used in one embodiment of the invention.

FIG. 53 is a vertical cross-sectional view of the power plant of thepresent invention, similar to the view first set forth in FIG. 1 above,but here showing the addition of an annular reaction turbine forcapturing the kinetic energy of the exhaust gases and generating shaftfor electrical power therefrom.

FIG. 54 is a cross-sectional view, taken across line 54—54 of FIG. 53,here showing stationary exhaust gas heat recovery section, the rotatingannular reaction turbine, aid the exhaust gas duct.

FIG. 55 is a partial isometric view of a second embodiment of thebiplane rotor of the present invention, similar to the view first setforth above in FIG. 5, here showing a solid metal matrix composite typeconstruction configuration.

FIG. A is a vertical cross-sectional view taken through line A—A of FIG.55, showing the construction of the solid type rotor.

FIG. B is a vertical cross-sectional view taken through line B—B of FIG.55, showing the construction of the solid type rotor, and also showingthe changing features of gap and fuel conduit diameter.

FIG. C is a vertical cross-sectional view taken through line C—C of FIG.55, similar to the view set forth in FIGS. A & B above, showing furthervariations in rotor dimensions with change in radial position.

FIG. D is a vertical cross-sectional view taken through line D—D of FIG.55, similar to the views in FIGS. A through C above, showing furthervariations in rotor dimensions with change in radial position.

FIG. E is a vertical cross-sectional view taken through line E—E of FIG.55, similar to the view set forth in FIGS. A through D above, showingfurther variations in rotor dimensions with change in radial position.

FIG. F is a vertical cross-sectional view taken through line F—F of FIG.55, similar to the views in FIGS. A through E above, showing furthervariations in rotor dimensions with change in radial position.

FIG. G is a vertical cross-sectional view taken through line G—G of FIG.55, similar to the views in FIGS. A through F above, showing furthervariations in rotor dimensions with change in radial position.

FIG. 56 provides an isometric view of an end cap for use with the solidtype rotor first illustrated in FIG. 55 above.

FIG. 57 illustrates a vertical cross-sectional view of the finished,operating position of the end cap just illustrated in FIG. 56, when thecap is affixed to the rotor.

FIG. 58 shows schematically the use of the power and heat generated inthe thrust module of the power plant for a variety of heat recovery;shaft work, or electrical co-generation activities.

FIG. 59 graphically shows a comparison of the general performancecharacteristics of the basic ramjet driven power generation plant, whencompared to gas turbines. Engine performance is shown in terms of heatrate.

FIG. 60 graphically shows the performance improvements available to thebasic ramjet driven power generation plant through (a) the addition ofheat recovery, and (b) by use of both heat recovery and a reactionturbine.

FIG. 61 graphically shows the cycle efficiencies, in terms of cycleefficiency, for various types of power plants, including the power plantof the instant invention.

FIG. 62 graphically shows the project cost of energy (cents per kilowatthour) for the basic ramjet driven power generation plant, when comparedto other types of power generation plants commonly in use today.

DETAILED DESCRIPTION OF INVENTION

The invention will be better understood and appreciated fromconsideration of a preferred embodiment thereof which, for purposes ofdescriptive clarity, includes simply a power plant with heat recoverytype exhaust gas cooling. It is of course appreciated that additionalfeatures and combinations with other power generation apparatus may bedesirable in particular circumstances. However, the power plant systemto be initially described below will be a basic building block in mostinstances of a power plant design due to the desirability of capturingthermal energy from combustion gases.

My power plant is based on high speed, supersonic propulsion phenomenonwhich allows the elimination of most moving parts which are common inother types of combustion power plants currently available.Simplification of the power generation apparatus allows initial capitalcosts to be minimized, and the superb system performance allowsoperating costs to be minimized.

Basic Power Plant

Referring now to the drawing, FIG. 1 depicts, in its operative powergeneration configuration a vertical cross-sectional view of a powerplant 100 constructed in accord with, and embodying, the principles ofthe present invention.

Key components of the power plant 100 include the following:

one or more thrust modules 102 a and/or 102 b suitable for oxidizing afuel supplied thereto from a fuel supply 103 and thus creating apropulsive thrust from the exhaust gases created;

a power output means such as a central rotating shaft portions 104 a and104 b;

a rotor 106 having one or more portions 106 a and/or 106 b (ideally, onerotor portion per thrust module) for rotatably connecting the thrustmodule(s) 102 a and 102 b with the output shaft portions 104 a and 104b.

Ideally, the thrust modules 102 a and 102 b are ramjet engines whichutilize oxygen from available airflow as an oxidant source.

In this embodiment, the shaft has an upper portion 104 a and a lowerportion 104 b for rotatably supporting rotor 106 and the appended thrustmodules 102 a and 102 b. The shaft portions 104 a and 104 b are hollowthus providing the necessary conduits 105 a and 105 b for flow of fuelto the thrust modules 102 a and 102 b from fuel supply 103.

Rotor 106 provides the means to rotatably connect and secure the shaftportions 104 a and 104 b to ramjet thrust modules 102 a and 102 b. Therotor 106 may include opposing upper biplane portions 108 a and 108 b,and opposing lower biplane portions 110 a and 110 b, as shown here, ormay be of an alternate configuration as further described hereinbelow.Opposing biplane rotor pairs 108 a and 110 a, and 108 b and 110 b, aresecured near their axial ends 112 a and 112 b to shaft portions 104 aand 104 b by hub means 114. The thrust modules 102 a and 102 b aresecured to the distal end 116 a of biplane rotors 108 a and 110 a, andto the distal end 116 b of rotor portions 108 b and 110 b.

Biplane rotor portions 108 a and 110 a, and 108 b and 110 b are shown aslaterally opposing portions (a “two spoke” configuration). However,other embodiments, such as a tri-rotor or quad-rotor (three or four“spokes,” respectively extending from a central hub) are also feasibleby use of the principles disclosed herein.

The basic rotating assembly, comprising the thrust modules 102 a and 102b, the upper rotor portions 108 a and 108 b, the lower rotor portions110 a and 110 b, hub means 114, and shaft portions 104 a and 104 b, arerotatably secured in an operating position by a support structure orhousing 120. Bearings 122 a, 122 b, 122 c, 122 d, 122 e, and 122 f, orsuitable variations thereof, provide adequate bearing support forrotation with minimum friction. The accompanying lubrication systems maybe provided by any convenient means by those knowledgeable in high speedrotating machinery, and need not be further discussed herein.

Support structure 120 includes several important features which areprovided to reduce aerodynamic drag on the rotors 108 a and 108 b, and110 a and 110 b. First, an upper housing portion 123 is provided with alower surface 124, and a lower housing portion 126 is provided with anupper surface 128. Surfaces 124 and 128, respectively, are located withminimal clearance between lower surface 124 and the upper biplaneportions 108 a and 108 b, and between the upper surface 128 and lowerbiplane rotor portions 110 a and 110 b, respectively. Thus, rotorportions 108 a, 108 b, 110 a, and 110 b may be rotated relative to thesupport portion 120, yet be securely held in a close fittingrelationship with the support portion 120 with minimum surface tosurface clearance in gap 129 a and 129 b.

A sweep air chamber 130 defined by wall 132 is provided as a conduit forair to flow past the rotor portions 108 a, 108 b, 110 a, and 110 b, sothat, for example, the rotor 108 b is not significantly affected by theaerodynamic wake of the just passed rotor 108 a. The air flow velocitynecessary to accomplish the desired objective will vary according to therotational speed of the rotor 106, and the radial length thereof, butmay be derived by the builder once other variables are identified.Adequate velocity of the air flowing through chamber 130 may be assuredby an induction fan (not shown) on the exit air stream, or othersuitable means.

As mentioned above, the upper high speed shaft 104 a and the lower highspeed shaft 104 b are hollow, thus including conduits 105 a and 105 b,respectively, to provide fuel from supply 103 to thrust modules 102 aand 102 b. From conduits 105 a and 105 b, fuel is routed through upperfuel passageways 140 a and 140 b and lower fuel passageways 142 a and142 b in rotor 106. The cross-sectional area of passageways 140 a, 140b, 140 c, and 140 d may be varied to accommodate the compression of fueldue to centrifugal effects.

The lower shaft portion 104 b also acts in conventional fashion totransmit mechanical power to the gear-box 150. The gear-box 150 reducesthe output shaft 104 b speed to a sufficiently low level to accommodatethe capabilities of the desired application. In FIG. 1, gear-box 150 isconnected by shaft 152 to primary electrical generator 154, suited togenerate electrical power for transmission to a power grid or otherelectrical load. However, shaft 152 could be applied directly to dodesired mechanical work.

Gear-box 150 is also shown connected by shaft 156 to starter motor 158.Starter motor 158 is supplied electrical power and control power from amotor control center. The starter motor 158 is configured to turn,through gear-box 150, the shaft 104 b so as to rotate thrust modules 102a and 102 b to a convenient tangential velocity so as to enable thestart of the ramjet engines. Once the ramjet engines of thrust modules102 a and 102 b are running, a properly designed starter motor 158 couldthen be shut down, and used in reverse as a generator of electricalpower.

Where appropriate hereinbelow, like numerals will be utilized toidentify like structures throughout the various figures, without furthercomment thereon.

Cogeneration

Exhaust 160 gases from thrust modules 102 a and 102 b may beconveniently cooled by an enthalpy extraction system 162 which surroundsand laterally encloses rotor 106. This system 162 includes a duct 164having therein hollow vanes 165, through which a secondary working fluidor coolant 166 is circulated. In the usual design, the working fluid 166will be water. The hot exhaust gases 160 from the thrust modules 102 aand 102 b flow through duct 164, impinging hollow vanes 165 and thusheating the fluid 166 therein. It may be convenient to design the system162 as a boiler so that the fluid 166 changes state, i.e., water becomessteam, as it is heated, and in such cases the stream indicated ascoolant out will be steam, suitable for use in heating, or in mechanicalapplications such as steam turbines.

An external support structure including legs 170 and 172 provide thenecessary structural support to enthalpy extraction system 162.Additionally, support structures 174 and 176 provide structural supportfor upper support structure 123 and interrelated components which housethe upper shaft portion 104 a.

For convenience, it may be desirable to locate legs 170 at grade 178level, and provide a utility vault 180 for containment of gear-box 150as well as generator 154 and starter 158.

Certain important features of the enthalpy extraction system 162 aremore clearly seen in FIG. 2, which is a horizontal cross section of thepower plant apparatus, taken through line 2—2 of FIG. 1. The lower rotorportions 110 a and 110 b of biplane rotor 106 are shown, joined by hub114. Upper surface 128 of lower housing support 125 can be seen belowrotor portions 110 a and 110 b. The location and configuration of theramjet thrust modules 102 a and 102 b can be clearly seen. Air flow duct130 provides upflow air supply.

The exhaust thrust vector 180 of the exhaust 160 from thrust module 102a, relative to the tangential direction 182 to the circumference ofrotation of rotor 106, may he outward by any predetermined angle alpha(α) which is convenient in the overall operational efficiency of theplant. The actual angle alpha (α) utilized is determined by the selectedlocation of the various heat transfer conduits, and by the use of areaction turbine (if any) as described herein below. The angle alpha (α)chosen is important since it helps to direct exhaust gases 160 moretoward the exhaust duct 164.

As illustrated, the coolant outlet conduits 184 are spaced radiallyoutward from a central axis 152 and are located at the periphery of theheat exchange section of duct 164; however, it will be appreciated thatthe coolant inlet and outlets may be varied as convenient for a giveninstallation and still accomplish desired heat exchange between asecondary fluid and the escaping exhaust gases. At the extremely highthrust velocities which will be encountered, an aerodynamicallyappropriate heat exchanger shape such as convex fins 165, having aleading edge 186 and a trailing edge 188 in the direction of exhaust gas160 flow, will help reduce backpressure on the heat exchange system 162.

Turning now to FIG. 3, an enlarged detail, similar to the view first setforth in FIG. 1, is provided. This enlarged drawing provides furtherdetail of the ramjet thrust module 102 a, the biplane rotor portions 108a and 110 a, central hub 114, rotating shaft portions 104 a and 104 b,and the heat transfer device 162 used for cooling exhaust gases 160.

No Heat Recovery Configuration

FIG. 4 is an enlarged detail of a power plant 200 similar to the onefirst illustrated in FIG. 3; however, in FIG. 4, the plant 200 does notinclude a heat transfer section for cooling exhaust gases. Supportstructures 170 and 172 provide support to exhaust duct forming members202, which in turn supports finned elements 204 and upper duct formingmember 206. The upper duct forming member 206 provides support to anupper support structure 174 and 176, just as in the earlier illustratedembodiments. One important feature first illustrated in this embodimentis the use of an exhaust duct 210 for containing therein the exhaustgases 160 for transport to an atmospheric release or further treatmentpoint.

Biplane Rotor Construction

The construction of one embodiment of the biplane rotor of the presentinvention is shown in FIG. 5 and FIG. 6.

FIG. 5 is a partial isometric view of a biplane rotor 220 of the typeprovided in the power plant apparatus 100 and 200 of FIGS. 1 through 4,respectively, above. In FIG. 5, a central hub structure 222 is shownhaving connected thereto upper 224 and lower 226 biplane portions. Thisdesign utilizes a biplane configuration which is carefully shaped tominimize aerodynamic drag of rotor 220; the design will be furtherdiscussed herein below. Each of the biplane portions 224 and 226 aregenerally triangular in chordwise cross-sectional shape. It can also beseen in FIGS. 5 and 6 that the upper 224 and lower 226 portions of thebiplane rotor 220 are separated by an air gap “G” which accommodates airflow therethrough. A ramjet type thrust module 230 is affixed to thedistal ends 232 and 234 of upper 224 and lower 226 biplane portions,respectively. Thus, it can be seen that the rotor 220 is a carefullydesigned low drag structural member which constrains the thrust modules230 to rotate with, and about the axis provided by, the high speed shaftportion 240.

The high tip speeds at which a rotor 220 must run in my power plantdesign in order to realize its superior performance levels necessarilyinduces very high stress levels in the rotor 220 and in the thrustmodule 230. As a result, rotor 220 stress levels represent a criticaldesign and operational problem. The thrust modules 230 and rotor 220 arebased on unique structural concepts and must be carefully designed toensure achievable and safe material stress margins.

In the embodiment illustrated in FIGS. 5 and 6, carbon fiber windings242 are located within the upper 224 and lower 226 biplane portions.Ideally, these carbon fiber windings 242 run continuously from tip totip (i.e., from the outer side of one thrust module to the outer side ofthe opposite thrust module) of the rotor portions 224 and 226, so as toform a high strength member to restrain the thrust modules 230, as wellas to reduce stress in other rotor materials to survivable levels. Toprovide sufficient rigidity to rotor portions 224 and 226, a metalmatrix composite material such as silicon carbide reinforced titaniummay be utilized to form the inlet 244 a and 244 b and outlet 246 a and246 b walls of the rotor, as well as centrally located verticalstructural members 248 a and 248 b. As shown, the upper leading edgeportion 252, upper trailing edge portion 254, and lower leading edgeportion 256 and lower trailing edge portion 258 each have a recessedportion denoted with suffix “r”. Likewise, the centrally locatedvertical members 248 a and 248 b have leading and trailing (noted hereand otherwise where appropriate with suffix “l” for leading or “t” fortrailing edge portions, respectively) portions with recesses therein(noted in the figures with the suffix “r” for recessed). The variousrecessed portions are configured to receive therein in a flush fittingfashion the respective upper protective covers 260 t and 260(l) andlower protective covers 262 t and 262(l). The protective covers 260 t,260(l), 262(t), and 262(l) are designed to provide an aerodynamicallysmooth upper 266 and lower 268 surface, while protecting the carbonfilaments 242 from oxidation. Protective end caps 270(l) and 270 t onthrust module 230 perform a similar function, and likewise fit theaforementioned recesses to the extent applicable. Each of protectivecovers 260 t, 260(l), 262 t, and 262(l) as well as protective caps270(l) and 270 t are securely affixed to the respective metal matrixcomposite portions, preferably by brazing or welding thereto so as toseal any seams between the various covers and the substrate rotorportion.

Hub portion 222 may be constructed of opposing sandwich portions 280 and282 (which are configured to accept therein the upper rotor 224 andlower rotor 226) and a central solid portion 284. Each of theaforementioned sections of the hub portion 222 may be constructed ofmaterials suitable for the anticipated structural loading at the designcentrifugal loadings. For the hub itself, conventional materials such ashigh strength steel may be sufficient in most applications.

Upper shaft portion 240 has therein an interior wall 288 which defines aconduit 290. The conduit 290 is used as a passageway for fuel to flow tothe thrust module 230. As the upper shaft portion 240 transitions tosandwich portion 280, conduit 290 turns from vertical to radial, and ispositioned near the center of central metal matrix compositestrengthening portions or “gutters” 248 a and 248 b. The cross-sectionalarea of conduits 290 may be varied as necessary to accommodate thecompressibility of the fuel being transported, so as to assure that fuelreaches thrust modules 230 at an adequate pressure.

Thrust Module Construction

The thrust modules 102 a, 102 b, and 230 shown above, and similarversions shown hereinafter, are critical components of my power plantdesign. Referring now to FIG. 7, a perspective view of the distal end ofa biplane rotor 220 is shown with a ramjet thrust module 230 attachedthereto.

FIG. 8 is a horizontal cross section, taken through line 8—8 of theramjet thrust module 230 of FIG. 7, looking downward at the constructionof the thrust module 230.

The thrust module(s) 230 is(are) the prime mover(s) of the instant powerplant invention. For a variety of reasons, it is convenient to constructthe thrust modules 230 as fixed geometry ramjets. The ramjet propulsioncycle and high rotor 220 tip speeds provide the thermodynamic basis forthe superior efficiency and performance of my power plant over prior artgas turbines, steam turbines and piston engines.

The ramjet 230 has five basic operational regions from front to rearalong the air/combustion gas flow path centerline 30, as follows:

1) the inlet 302, through which air is admitted to the thrust module 230and in which the velocity of the incoming air stream 304 is reduced asram air pressure is developed;

2) the transition section 306, where the air flow slows and reaches mach1.0 (M=1);

3) the combustor 308, which includes a flame holder 310 (fuel isintroduced into the combustion zone and hot combustion gases arereleased from the combustion zone);

4) the throat 312, where the exit exhaust gas flow is choked; and

5) the nozzle 314, through which combustion gases 316 are ejectedrearward at high velocity.

Construction of the thrust module 230 may be better understood byreviewing a series of cross-sectional views taken along the length ofthe module 230.

FIG. 9 is a vertical cross-sectional view, looking rearward in thedirection of the exhaust 316 in the thrust module 230 of FIG. 7, takenat the station indicated by line 9—9 in FIG. 8. This view shows theleading edge 320 of thrust module 230, and the minimum cross-sectionalarea of the interior air flow passageway 321, defined by an innermostinterior surface 322 of sloping transition section 306 wall 324, and bythe outermost surface 326 of the inlet ramp of thrust module 230.

FIG. 10 is a vertical cross-sectional view, looking rearward in thedirection of the exhaust 316, cut through the thrust module 230 of FIG.7, taken at the station indicated by line 10—10 of FIG. 8. This viewshows the thickening inlet wall portions 320 and 332 of the thrustmodule 230, as well as the air flow passageway 321 already seen in FIG.9 above.

FIG. 11 is a vertical cross-sectional view, looking rearward in thedirection of the exhaust 316, cut through the thrust module 230 of FIG.7, taken at the station indicated by line 11—11 of FIG. 8. This viewshows the outer cap 270(l) of the thrust module 230, as well as thefirst layer 336 of the reinforcing carbon fiber windings 242 which wraparound the end of the thrust module 230.

FIGS. 12 through 15 are vertical cross-sectional views, looking rearwardin the direction of the exhaust, cut through the thrust module 230 ofFIG. 7, taken at the stations indicated by reference of FIG. 8, similarto FIGS. 9 through 11 above. FIGS. 12 through 15 show the varyingthickness of the reinforcing carbon fiber windings 242 (carbon fiberlayers 336, 338, 340, 342, and 344) as well as the shape of the interiorof the thrust module air flow path 321.

FIG. 16 is a vertical cross-sectional view, looking rearward in thedirection of the exhaust, cut through the thrust module 230 of FIG. 7,taken at the station indicated by line 16—16 of FIG. 8. This view showsthe outer cap 270 t of the thrust module 230, as well as the shape ofthe interior of the thrust module 230 air flow path 321 at this point inthe exhaust section.

FIG. 17 is a vertical view, looking forward toward air inlet 304 fromthe rear of the thrust module 230 of FIG. 7, taken at the stationindicated by line 17—17 of FIG. 8.

Startup of Ramlet

For startup, an auxiliary power system is used to accelerate the rotor220 and thrust module 230 to a sufficiently high rotating speed so thatthe ramjet operation of thrust module 230 can be initiated. Attention isagain referred back to FIG. 1, where it can be appreciated that therotating components are designed with sufficient strength to allow thestarter motor 158 to accelerate the rotor 106 and thrust modules 102 aand 102 b up to a sufficient speed so as to support ramjet operation.The required airspeed of thrust modules to begin ramjet operation willvary widely depending upon a specific design, however airspeeds in themore narrowly defined range of mach 1.5 to 2.0 might be expected toprovide adequate starting behavior for the ramjet configurationsdescribed herein.

After the thrust module(s) 102 a and 102 b begin to generate sufficientthrust, the starter motor 158 can be switched to a power generating modeof operation, and generate power along with the primary generator.

The power plant system requires a fuel control valve 350 to adjust thefuel to air mixture, as this ratio varies with both the thrust module102 a and/or 102 b tip speed and with desired system output powerlevels. In my power plant design, the entering fuel is compressed bycentrifugal forces as it flows through passages 140 a, 140 b, 142 a, and142 b in rotor 106 outward toward the thrust modules 230. This isparticularly important where a gaseous fuel such as methane is utilized.

As a result of the compression of fuel, and due to the compression ofincoming air, the startup of the thrust modules 102 a and 102 b must becarefully attended to by the designer. Several options for accomplishingthis task are addressed in FIGS. 18 through 24.

FIG. 18 is a horizontal cross-sectional view, similar to the view firstset forth in FIG. 8 above, showing a first alternate configuration forthe interior of a ramjet thrust module 360, utilizing a reverse Lavalinternal contraction type inlet. Note in particular the shape of theinlet 362 and transition 364 surfaces FIG. 18 also shows the outermostareas requiring carbon fiber 366 or similar reinforcement for safeoperation at normally encountered centrifugal loads.

FIG. 19 is a horizontal cross-sectional view, similar to the view firstset forth in FIG. 8 above, showing a second alternate configuration fora ramjet thrust module 370, utilizing a mixed contraction type inlet,wherein the interior leading edge 372 creates a shock wave 374 whichexactly impinges upon the exterior leading edge 376 so as to contain thereflected shock 377 within the inlet area. FIG. 19 also shows areasrequiring carbon fiber 378 reinforcement for operation in the presentinvention.

In FIG. 20 a detailed horizontal cross-sectional view, similar to theview first set forth in FIG. 8 above, shows a third alternateconfiguration for the interior of a ramjet thrust module 380, utilizingan ejector augmented flow path. Here, an ejectant 382 may be supplied toaugment the fluid flow through the combustor section 384 of the ramjet380. In some cases, the ejectant 382 may be necessary to induce airflowto start through the ramjet 380. The FIG. 20 also shows areas requiringcarbon fiber 386 reinforcement for operation in the present invention.

FIGS. 21A, 21B, and 21C illustrate air flow spillage and shock wavelocation for the startup of a mixed contraction inlet ramjet thrustmodule 390. The mixed contraction inlet ramjet 390 is similar to thesecond alternate thrust module 370 configuration first illustrated inFIG. 19 above. These FIGS. 21A, 21B, and 21C also show areas requiringcarbon fiber 392 reinforcement for operation in the present invention.

FIG. 21A shows shock wave location and spillage for operation of aramjet thrust module 390 at an airspeed well below design mach number.The FIG. 21A also shows areas requiring carbon fiber 392 reinforcementfor operation of the present invention.

FIG. 21B shows shock wave location and spillage for operation of a mixedcontraction inlet ramjet 390 slightly below design mach number. Thefigure also shows areas requiring carbon fiber 292 reinforcement foroperation in the present invention.

FIG. 21C shows the shock wave location and the captured airstream tubeas would be present in the operation at design mach number of a ramjetengine 300 having a mixed contraction inlet.

FIG. 22 illustrates the airflow configuration for an internalcontraction inlet ramjet 400. FIG. 22A illustrates a generalized crosssection configuration of an internal contraction type ramjet thrustmodule 400, taken across section 22A—22A of FIG. 22; it is similar tothat first illustrated in FIG. 8 above. FIG. 22A also shows areasrequiring carbon fiber 402 or other appropriate reinforcement foroperation in the present invention. Note that an imaginary line drawnbetween the leading edges of interior 404 and exterior 406 wall inletsform a plane perpendicular to the free stream airflow. The inflow airstream is compressed up an inlet ramp 408, and as will be seen in FIGS.33 and 34 below, inlet shocks are captured well inside the ramjet 400.

FIG. 23 illustrates the airflow configuration for a self-starting, mixedcompression, inlet ramjet thrust module 410. FIG. 23A illustrates ageneralized cross section configuration of a mixed compression typeramjet thrust module 410, taken across section 23A—23A of FIG. 23. Thisis an alternate configuration to the type of ramjet design firstillustrated in FIG. 8 above. As before, use of carbon fiber windings 412or other high strength techniques are required to provide adequatestructural strength to withstand the forces encountered at highrotational speeds. In the mixed compression inlet ramjet 410, a rakeangle Y is provided so that the lip 411 of the interior inlet wall 412and the exterior inlet wall 414 are offset by the angle Y so that theshock caused by the inlet wall 412 is captured by lip 414 of theexterior wall 416 when the ramjet 410 is operating at the design Machnumber.

FIG. 24 shows a generalized cross section configuration of an internalcompression type ramjet thrust module 420, similar to that firstillustrated in FIG. 8 above, now showing the combustor 422 location inthe thrust module, as well as describing several key regions includingnozzle 424. For reasons set forth below, this type of configuration ofthrust module has certain advantanges for operation of the presentinvention.

To summarize the discussion of FIGS. 18 through 24, two basic classes ofinlets have been introduced: (a) internal compression inlets, and (b)mixed compression inlets. It is well established that, in general,optimal internal contraction inlets will not start at their design Machnumber. In order to establish the desired internal shock structure, theinflow must either be accelerated to a Mach number greater than thedesign Mach number and then reduced after starting to the design Machnumber, or the throat area must be temporarily increased to “swallow”the shock structure and thus induce startup. Depending upon thecontraction ratio and Mach number, it may be impossible to increase theinflow Mach number to a sufficiently high level so as to start theinlet.

Also, due to the centrifugal loads associated with full speed operation,although it might be possible with some difficulty, it is undesirable toprovide a variable geometry mechanism to provide the increased areanecessary for startup, as such a mechanism would add additional weightand use up valuable space. In the absence of a workable variablegeometry mechanism and without the ability to sufficiently overspeed theinlet so as to start it, ejector augmention would be required to “start”an internal contraction inlet.

Alternatively, a mixed compression inlet could be designed which would“self start.” The mixed compression type inlet, as shown in FIGS. 21A,21B, 21C, 23, and 23A, does not require overspeed or a variable geometrythroat to start. It can be designed so that the shock structureprogressively reaches its desired position as indicated in FIGS. 21A,212B, and 21C and is fully started when the inlet reaches its designmach number.

Due to the absence of strong shocks in properly designed internalcompression inlets, internal compression inlets are generally moreefficient than mixed compression inlets. However, the difficultyassociated with starting the internal contraction inlet makes the“self-starting” mixed compression inlet a highly desirable alternateembodiment.

Theoretical Basis of Ramjet Design

With the foregoing general description of the apparatus and method ofoperation of several embodiments of a power plant system serving to setforth the basic elements of the present invention, before otherembodiments and variations are described, it will be useful to considerto an appreciable extent the theoretical analysis of the instant system.Accordingly, the following analysis is offered by way of explanation andis not intended to expressly or impliedly limit the scope of theinvention.

There are several fundamental factors which are important to theover-all thrust module design and operation. Such factors include: (a)mach number, (b) power output variability, (c) fuel type, and (d)maximum allowable combustion temperatures.

In the design of the thrust module, the Mach number must be selected,and the design must then include sufficient structural and materialselection and tolerances to allow for acceleration to and operation atthe desired operational velocity. The Mach number is commonly determinedby the following equation: $\begin{matrix}{M = \frac{v}{a}} & (1)\end{matrix}$

The required variations in thrust module output, and the resultant plantsystem power output, must be understood and accommodated.

The fuel type to be used must be determined, and the fuel feed systemmust accommodate the fuel type selected. Additional factors to beconsidered in fuel system design are compressibility, temperature,corrosion or erosion tendencies, and similar fluid flow phenomenon of aspecific fuel type.

The fuel type selection will also in large part determine the combustiontemperatures, and thus dictate the required materials or influencestructural requirements to accommodate the anticipated exhaust gastemperatures.

As noted above, in its simplest form, the thrust module is a fixedgeometry ramjet. The operation of such a ramjet engine 500 is depictedschematically in FIG. 25. In order to easily understand the ramjet 500operation, it is convenient to assume that the ramjet 500 is stationaryand that an airstream 502 flows toward the ramjet at velocity v₀. Then,consider that the approaching stream of air is of sufficiently largecross section so that the pressure is atmospheric along the boundariesof the control volume. (See FIG. 25.)

The air flow around the outside of the thrust module 500 suffersmomentum losses due to skin friction, so that the mean velocity v₇ ofthe external air (not in exhaust gas stream) at station 6 is less thanv₀. This momentum loss constitutes the viscous drag on the exterior ofthe thrust module. Such viscous loss cannot be avoided without the useof complicated boundary layer bleed orifices on the external surface ofthe module. However, since such bleed orifices are presently complex andexpensive, in view of the fact that such viscous losses are reasonablylow, it is unnecessary to include a boundary layer control system toaccomplish acceptable baseline power plant operational efficiencies.

Potential pressure drag, due to a change in a cross-sectional area or achange in a local pressure field, is of greater concern. At theoperational speeds of the present power plant, the effect of such anarea or pressure variation is aerodynamic drag. Pressure drag can easilyexceed the above discussed viscous drag by several orders of magnitude.Thus, avoidance of pressure drag is quite important. Therefore, mythrust module 500 has been developed to minimize pressure drag byconstructing the thrust module 500 of a constant externalcross-sectional shape (i.e., the shape and size is repeated whensequentially examined in cross section perpendicular to the axis of flow(spanwise) from a forward cross section to a rearward cross section).This construction technique is apparent by examination of the crosssections shown in FIGS. 8 through 17 above.

The ramjet inlet section captures and compresses an impinging inlet airstream. The compressed air stream thus provides the oxidant for mixingwith a fuel which is supplied to the ramjet thrust modules 500 in theform of a convenient fuel source such as natural gas (consisting ofessentially methane). The fuel is oxidized in the thrust module(s) toproduce combustion gases. The gases expand, and the exhaust gas flowescapes at high velocity v₅ to create thrust. This exhaust gas flowvelocity changes to v₆ when pressure equilibrium with the atmosphere isestablished.

From FIG. 25 it is apparent that the thrust of the engine is equal tothe difference between the momenta of the gases passing throughreference stations 6 and 0. Thus

F=({dot over (m)}₀−{dot over (m)}₁)v₇+({dot over (m)}₁+{dot over(m)}_(f))v₆−{dot over (m)}₀v₀  (2)

or

F={dot over (m)}₁(v₆−v₀)+{dot over (m)}_(f)v₆−({dot over (m)}₀−{dot over(m)}₁)(v₀−v₇)  (3)

The last term in equation (3) represents the momentum loss due toexternal drag as discussed above. This term can become prohibitivelylarge if due care is not exercised in configuring the thrust moduleportion of the system.

In the ideal case, the inlet and exhaust are expanded so that thepressures are ambient; then

A₀=A₁  (4)

and

A₅=A₆  (5)

Therefore, the net thrust of the thrust module is

F={dot over (m)}₁(v₆−v₀)+{dot over (m)}_(f)v₆  (6)

However, as a practical matter, equations (4) and (5) do not alwayshold, and it is found that the equation for thrust should include a terminvolving the difference in pressure between the inlet and exit. In suchcases,

F={dot over (m)}₁(v₆−v₁)+{dot over (m)}_(f)v₆+(p₆A₆−p₁A₁)  (7)

However, the current embodiment of the thrust module can be adequatelydeveloped on the basis that equations (4) and (5) both apply. Thesimilarity of equation (7) to the expression for rocket thrust will beapparent to those skilled in the art of high speed and space propulsionsystems. While such a rocket type thrust module will achieve many of theobjectives of the present invention, and my power plant concept isdirectly applicable thereto, due to the need in a rocket to makeprovisions for external supply of oxidant, as well as increased dragexperienced in such a design, the currently preferred embodiment isconsidered to be a ramjet thrust module 500.

Ramjet thrust calculations are considerably more complicated than arethe calculations for rockets. This is because in ramjets, the exhaustgeometry and gas velocity depend upon the interaction and balancebetween the pressure developed in the ramjet inlet and the pressuredeveloped in the ramjet combustor.

In a ramjet, the stream thrust T at a particular cross section in theramjet flow path is defined by the equation.

T=pA+{dot over (m)}v=pA(1+γM²)  (8)

The stream thrust T is a particularly useful quantity in ramjetcalculations because the difference in stream thrust between twostations is equal to the thrust exerted in an axial direction on theduct walls between the inflow and outflow planes. The stream thrust Tmay also be expressed as a function of mass flow, stagnation temperatureand Mach number. Thus $\begin{matrix}{{T = {{\overset{.}{m}\left( \sqrt{\frac{2\left( {\gamma + 1} \right)}{\gamma}R_{g}T_{t}} \right)}{\varphi (M)}}}{where}} & (9) \\{{\varphi (M)} = \frac{2 + {\gamma \quad M^{2}}}{\sqrt{2\left( {\gamma + 1} \right){M^{2}\left( {1 + {\frac{\left( {\gamma - 1} \right)}{2}M^{2}}} \right)}}}} & (10)\end{matrix}$

The Mach number goes to unity (M=1) at the transition section (station2) and at the throat (station 4) of the nozzle.

Another method for determining stream thrust is to begin by defining afuel specific impulse, S_(f), and an air specific impulse, S_(a), by therelations $\begin{matrix}{{{\overset{.}{w}}_{f}S_{f}} = {{{\overset{.}{w}}_{a}S_{a}} = {\overset{.}{m}\sqrt{\frac{2\left( {\gamma + 1} \right)}{\gamma}R_{g}T_{t}}}}} & (11)\end{matrix}$

With the quantities defined above and the equation of state,determination of stream thrust is possible. Thereafter, ramjet designcalculations are straightforward but involve successive approximations.A procedure is summarized below to illustrate the method.

Using the engine configuration shown in FIG. 8 and assuming the designcondition of an oblique shock system between stations 1 and 2, we maydesign an engine as set forth below.

With sonic conditions existing at the throat to the exit nozzle, it ismost convenient to express the internal losses in terms of totalpressure ratios, so that $\begin{matrix}{p_{t5} = {p_{t0} \times \frac{p_{t1}}{p_{t0}} \times \frac{p_{t2}}{p_{t1}} \times \frac{p_{t4}}{p_{t2}} \times \frac{p_{t5}}{p_{t4}}}} & (12)\end{matrix}$

The free stream stagnation pressure (P_(t0)) is determined by the thrustmodule speed. The total pressure at the inlet inflow plane (P_(t1)) isgenerally equivalent to the free stream stagnant pressure (P_(t0)).

The inlet efficiency is the ratio of the total pressure at the outflowplane of the inlet (P_(t2)) divided by the total pressure at the inflowplane of the inlet (P_(t1)). The ratio of the total pressure at thethroat of the nozzle (P_(t4)) to the total pressure at the outflow planeof the inlet (P_(t2)) defines the efficiency of the flow field betweenthe inlet and the nozzle, and includes losses across the combustor 522due to fuel injection, drag and heat addition. Those knowledgeable inthe art may also estimate the efficiency from past experience with othersystems. The total pressure ratio across the exit nozzle is the totalpressure at the outflow plane of the nozzle (P_(t5)) divided by thetotal pressure at the inflow plane of the nozzle (P_(t4)).

In order to determine the net thrust of the thrust module, thedifference between the stream thrust at stations 1 and 5 must bedetermined. Assuming that the nozzle exit is sized to provide anexpansion to ambient pressure (i.e. p₅=p₀) , the exit Mach number atstation 5 (M₅) may be determined from the ratio of static pressure atstation 5 to total pressure at station 5 (p₅/p_(t5)). The exit streamthrust may then be expressed as:

T₅=p₀A₅(1+γ₅M₅ ²)  (13)

The exit stream thrust (thrust at station 5) may also be expressed:

T₅={dot over (w)}_(a)S_(a)φ(M₅)  (14)

The air specific impulse (S_(a)) is a function of the inlet totaltemperature, the fuel, the fuel-to-air ratio, and the combustionefficiency. The inlet stream thrust may be expressed as:

T₁=p₀A₁(1+γ₀M₀ ²)  (15)

Defining the mass flow at the inlet, $\begin{matrix}{{f\left( M_{0} \right)} = \frac{{\overset{.}{w}}_{a}\sqrt{T_{t0}}}{p_{0}A_{1}}} & (16)\end{matrix}$

More usefully, the inlet stream thrust may also be expressed as:$\begin{matrix}{T_{1} = {\frac{{\overset{.}{w}}_{a}\sqrt{T_{t0}}}{f\left( M_{0} \right)}\left( {1 + {\gamma_{0}M_{0}^{2}}} \right)}} & (17)\end{matrix}$

Therefore, the net propulsive thrust of the ramjet is the differencebetween the stream thrust at stations 1 and 5, such that:$\begin{matrix}{{T_{5} - T_{1}} = {{{\overset{.}{w}}_{a}S_{a}{\varphi \left( M_{5} \right)}} - {\frac{{\overset{.}{w}}_{a}\sqrt{T_{t0}}}{f\left( M_{0} \right)}\left( {1 + {\gamma_{0}M_{0}^{2}}} \right)}}} & (18)\end{matrix}$

The net thrust is thus expressed in terms of the air flow captured bythe inlet. The required inlet and exit areas may be determined fromrelations (13) and (15).

The combustion chamber cross-sectional area A₂ and the combustionchamber cross-sectional area A₄ is determined by the allowable pressurelosses across the combustor. Because these pressure losses are excessiveat high Mach numbers, the cross-sectional areas A₂ and A₄ must beselected to maintain low velocities, thus resulting in low pressurelosses. Consequently, the most desirable values for M₂ (Mach number atstation 2) may be selected from approximately 0.15 to 0.50; however, itmay be possible to operate outside this range, normally with somewhatreduced efficiency.

FIG. 26 shows, in graphical form, the variation in thrust output fromthe thrust module at various throttle settings, for a design Mach numberof 3.5. FIG. 27 shows in graphical form the variation in thrust modulethrust at various Mach numbers. FIGS. 26 and 27 are based on a set ofspecific assumptions regarding thrust module sizes, free streamconditions, and fuel source. In both cases, a thrust module as indicatedin FIG. 8 with an inlet cross-sectional area of 0.087 ft² is assumed. Atthe in flow Mach number indicated in FIG. 26, this results in a massflow of air into the inlet of approximately 25.3 pounds mass per second.Results are shown for combustion of natural gas. In the case of FIG. 27,the mass inflow varies with inflow Mach number. In both cases a freestream temperature of 549° R. and a pressure of 2116 pounds per squarefoot is assumed. Component performance levels consistent with wellestablished test data are assumed for the inlet, transition section,combustor, and nozzle.

Because the thrust module thrust determines the overall power plantoutput, the thrust from the module is an important figure of merit forthe thrust module and overall plant output levels. The thrust modulethrust levels and the overall plant output levels increase in directproportion with the mass captured and processed by the thrust module.Thus, doubling the inlet area and mass capture results in doubling thethrust generated by the thrust module, and thus results in doubling thepower output of the system.

Thrust Module Performance

Using the basic performance equations set forth above, a thrust moduledesign has been developed which yields excellent performancecharacteristics. The geometry of the flowpath, including its shape andarea variations, is the basis for the performance characteristics of thesystem.

For structural reasons the basic flowpath of the thrust module is at allstations semi-circular. This basic configuration is represented in FIGS.9 through 17. The radially innermost surface is planar. The radiallyoutermost surface is semi-circular in radial cross section with acircumferential area variation of the character indicated in FIG. 8.

Based on a given inlet contraction ratio, combustor shape and expansionratio, the basic thrust module performance variation with tip Machnumber is shown on FIG. 27. It is important to remember that FIG. 27does not represent the overall engine performance. As already discussed,the thrust produced by the thrust module drives the rotor which in turndrives the shaft which is used to produce useable mechanical work.Integrated power plant performance will be discussed below.

FIG. 26 shows the variation in the thrust module output over a range ofthrottle settings. FIG. 26 also shows the variation in combustortemperature with varying throttle settings. The combustor temperature isa critical factor. Combustor temperature must be balanced with inflowrate and thrust module materials so as to maintain structural integrityin the combustor walls.

The thrust module is made of a material with desireable high temperaturecapability covered with an oxidation/wear protection coating. Candidatematerials include hot isostatic pressed alumina, silicon nitride,zirconia, beryllia, and silicon carbide.

Rotor Aerodynamic Design

As already discussed, a key feature of the instant power plant is therotor (e.g., rotor 220 above) which attaches the thrust module to thehigh speed shaft. This rotor is rigidly attached to and rotates with thethrust module. Two design parameters of the rotor are extremelyimportant. First, the rotor must be constructed of materials whichenable it to survive the extremely high centrifugal loads encounteredwhile the thrust module is rotating at a tip speed in the Mach 3.5range, i.e., the rotor must be capable of withstanding extremely hightensile stress. Second, at such speeds, minimizing the rotor's overallaerodynamic drag is critical.

Since the rotor moves with the thrust module, the speed with which itmoves through the air varies along its length, proportional to theoperational radius of any position along its length. Basically, thelocal tangential velocity at any radius outward along the rotor willvary from zero at the axis of rotation to Mach 3.5 or more at the tip.Thus, the air flow over the rotor varies from subsonic through transonicand up into supersonic speeds.

The aerodynamic drag experienced by structures of various shapes movingat supersonic speeds in the Mach 3 to 4 range can become extreme. Sinceoverall power plant efficiency decreases with increasing rotor draglosses, it is self evident that the rotor drag losses must be limited toa small fraction of the thrust generated by the thrust modules.Obviously, if the parasitic system power losses due to rotor drag becomeappreciable, the efficiency of the instant power plant would dropdramatically, potentially to levels below that of conventionalgeneration systems.

Several different rotor configurations have been considered to providethe required strength at minimal aerodynamic drag. In one embodiment, asillustrated in FIG. 5 above, the rotor 220 includes a pair of generallytriangular shaped arms (224 and 226) extending from the rotor huboutward to the thrust module 230. In another embodiment, as will be setforth hereinbelow, the rotor is provided as a continuous disc, with thethrust modules mounted at the rim of the disc. In yet anotherembodiment, also set forth hereinbelow, these two concepts are combined,with a pair of arms extending outward from a central disk at two or morebalanced locations, with thrust modules located at the distal end ofeach pair of arms. The flow fields around these various embodiments arefundamentally different. Therefore, the theoretical analysis of eachwill be discussed separately below.

The drag of a rotor having a discrete pair of arms or blades can beaccurately calculated by analyzing the airflow over the various rotorcross-sectional shapes. Due to structural and aerodynamicconsiderations, it is preferable that the cross section of the rotorvary along its span. The overall drag on the rotor can be determined byadding up the drag contributions from the various cross sections whichexist over the span of the rotor. One specific embodiment of thisconcept is illustrated in Table I and will be discussed further inconjunction with FIG. 35 below.

In view of the importance of rotor drag reduction, it will be useful tobriefly consider the theoretical basis for analysis of supersonic flow,before other embodiments and variations are described. In order to allowthe reader to better appreciate the importance of the shape of therecommended rotor designs, several aerodynamically acceptable shapes andseveral alternative but aerodynamically unacceptable shapes (which mightbe structurally useful) for rotors will be explored. In supersonic flow,pressure or wave drag exists even in an idealized, non-viscous fluid.This supersonic drag is fundamentally different from the friction dragand the separation drag that are associated with the boundary layers ina viscous fluid. The latter are easily calculated by those trained inthe art and are in any event of considerably lesser importance than thesupersonic pressure drag at the desired rotor velocities of interest forefficient operation of my power plant. Accordingly, the followinganalysis is offered by way of explanation and is not intended toexpressly or impliedly limit the scope of the invention.

Reference is now made to FIG. 28, wherein a diamond shaped airfoil 600is illustrated. In FIG. 28, the free stream at velocity M₁ in region 1has a static pressure of P₁, as noted in the pressure distributiondiagram at the bottom of FIG. 28. The nose shock 602 compresses the airflow to the positive pressure P₂ in region 2, and the centered expansionfan 603 at the shoulder 604 expands pressure to a negative pressure P₃in the region 3. It is important to note that at trailing edge 606, theshock 608 recompresses the pressure in region 4 to P₄, which isessentially equal to the free stream value, P₁. The pressure P₂ and P₃both retard the progress of the diamond shape 600 through the airstream.Thus, the diamond shape 600 is aerodynamically unacceptable for a rotordesign because the drag component, computed by integrating the pressureover the projected area, is unacceptably large.

The drag “D” on the airfoil 600, due to the overpressure on the forwardface 610 and underpressure on the rearward face 612, may be expressed,for a unit span, as:

D=(p₂−p₃)t  (19)

where t is the thickness of the section at the shoulder. The values ofP₂ and P₃ are easily found from shock charts and tables of thePrandtl-Meyer function, as might be found in any convenient aerodynamictextbook.

In FIG. 29, another alternative shape is illustrated. Here, a curved,bi-convex airfoil section 620 is provided. The airfoil section 620 has acontinuous decrease in pressure as the air flow expands along the uppersurface 622 (and lower surface 623) as seen in the pressure distributiondiagram at the bottom of FIG. 29. For the leading edge shock 628 to beattached, it is necessary that the nose 624 be wedge shaped. In the caseof a half angle greater than the critical angle, the shock would becomedetached, as indicated by the broken line shock location 628′.

As with the diamond shape 600 illustrated above, the convex shape 620has an unacceptably high pressure drag. At the leading edge or nose 624,the pressure increases to P₂ from the free stream pressure P₁. Thepressure drops across the body, changes from positive to negative at thetangential point 627, and reaches P₃ at the trailing edge 626. Again,the integral of pressure over the projected area results in excessivedrag.

In FIG. 30, a flat plate 630 with an angle of attack beta (β) is shown.Like the diamond 600 and convex 620 shapes, there is no upstreaminfluence on the airflow due to the presence of the plate 630, so theair stream line 632 ahead of the leading edge 634 is straight, and is atan airspeed of M₁. The portion of flow over the upper side 636 is turneddownward through an expansion angle H by means of a centered expansionfan 638 at the leading edge 634, whereas on the lower side 640 the flowis turned downward through a compression angle H (equal to (β)) by meansof an oblique shock 642. From the negative pressure P₂ on upper side ofthe plate 630 and the positive pressure P₂′ on the lower side of theplate 630, the lift “L” and drag “D” are computed very simply. They are,

 L=(p′₂−p₂)c COS β  (20)

D=(p′₂−p₂)c sin β  (21)

where c is the chord. Unfortunately, the drag experienced by flat plate630 is unacceptably high for operation at the supersonic speeds where myramjet power plant is most efficient.

As indicated in FIGS. 28, and 29 above, both the diamond shaped airfoil600 and she bi-convex airfoil 620 experience large pressure or wave dragdue to the surface pressure distributions induced by the presence of theshock waves. If this pressure drag existed on a rotor, its effect wouldbe magnified by the local rotor speed. The power required to overcomesuch drag would be described as follows: $\begin{matrix}{P_{ower} = {\int_{hub}^{tip}{r\quad w\quad D{r}}}} & (22)\end{matrix}$

Because the local rotor velocities are very large, small drag values canconsume substantial power levels and thus result in economicallyprohibitive reductions in system efficiencies.

So far no mention has been made in these examples of the interactionbetween the shock waves and the expansion waves. This is because in theshapes illustrated in the FIGS. 28 through 30 above, interaction betweenshock waves and expansion waves is not of assistance in the reduction ofdrag, as it is in the biplane configuration as further illustrated andexplained in conjunction with FIGS. 33 and 34 below.

To examine the interaction of the shock waves and the expansion waves,it is necessary to examine a larger portion of the flow field than wasset forth in the FIGS. 28 and 30 above. Thus, FIGS. 31 and 32 showsimilar versions of two examples set forth in aforementioned previousFIGS.

The expansion fans 650 (in FIG. 31) and 652 (in FIG. 32) attenuate theoblique shocks 654 and 656, respectively, making them weak and curved.At large distances from the leading edges 658 and 660, the shocks 654and 656, respectively, approach asymptotically the free stream Machlines.

The reflected waves are not shown in FIGS. 31 because in shock expansiontheory, the reflected waves are normally neglected. Their effect issmall, but in an exact analysis they would have to be considered. In thevarious shapes, the wave system extends to very large distances from theshape; at such distances all such disturbances are reduced toinfinitesimal strengths. For a diamond shaped cross section and for alifting flat plate (unlike the case for the biplane shape discussedbelow) the reflected waves do not intercept the airfoil at all, andhence do not affect the shock-expansion result for the pressuredistribution. That is important because it is the pressure distributionwhich is of primary concern in evaluating the supersonic wave drag.

The previous discussion focussed on the characteristics of supersonicflow. This focus is necessary because the portion of the rotor whichcontributes the vast majority of drag is the outboard portion. Thatportion of the rotor would be supersonic during normal thrust moduleoperation. Thus, the dominating aerodynamic effects which contribute tooverall rotor aerodynamic drag are embodied by the relatively simpleshock expansion characteristics just discussed. While the inboard andtransonic regions of the rotor may have appreciable drag levels, thefact that those regions of the rotor are moving at a lower velocitymeans that the power consumed in overcoming the drag from those regionsis smaller than that required with respect to the outboard, supersonicportions of the rotor.

A wealth of empirical correlations and analytical methods exist to showthe subsonic and transonic flow over various possible rotorcross-sectional shapes. Such methods may be fully developed for rigorousanalysis by those skilled in the art. However, since the contribution ofsuch subsonic and transonic drag is appreciably smaller than from thesupersonic drag on the rotor, as just discussed above, analysis of suchregions is not critical to a basic understanding of the principles ofthe present invention.

Biplane Rotor Aerodynamic Performance In one embodiment, my power plantincludes a biplane rotor design which minimizes aerodynamic drag lossesthrough use of geometric shaping which provides shock cancellation toeliminate pressure drag. From the above discussion of supersonic flowover various shapes, it is clear that shock waves are formed whereverthere is a deflection of the high speed airflow by a shaped surface.Recognizing this phenomenon, in order to use it to advantage rather thanto merely avoid the disadvantages, I have designed the outboard regionsof the rotor (e.g., rotor 220 above) utilizing a configuration comprisedof two triangular cross sections or biplane portions. Each of thesetriangular portions is carefully contoured to result in shockcancellation within the space between the biplanes. The important resultis that a biplane rotor can be supplied having essentially constantpressure within the region between the leading edge and trailing edge ofthe biplanes at the design Mach number. Such a configuration virtuallyeliminates supersonic pressure drag on the rotor.

To illustrate this supersonic pressure drag reduction technique by shockcancellation in the biplane structure of the present invention (incontrast to drag arising in the previously discussed FIGS. 28 through32), a typical biplane rotor 660 cross section is shown in FIG. 33. Theleading shock structure 662 a resulting from passage of the biplanethrough the airflow at speed M_(1a) is indicated. These shocks 662 a arecancelled by the expansion created by interior upper shoulder 668 u andlower shoulder 668(l) of the biplane. As a result, the pressuredistribution is even (at pressure=P₂) throughout the interior of thebiplane 660. Because the outside surfaces (upper surface 664 and lowersurface 666) are flat, there is no airflow deflection from the outersurfaces, and therefore no exterior shock waves are formed.

The significance of the cancellation of shock 662 a by use of thebiplane rotor technique is clear when one evaluates the pressuredistribution which results from the shock 662 a at design condition,where M_(1a)=Design airspeed (see the pressure distribution diagramportion at the bottom of FIG. 33.). FIG. 33 shows the pressuredistribution induced by the shocks on an ideally designed and operatedbiplane structure 660. At the leading edge 669, the pressure is raisedfrom the free stream pressure P₁ to the interior pressure P₂. Theinterior pressure remains essentially constant at pressure P₂ until thetrailing edge 670 is reached, i.e., P₂=P₃.

It can be seen from the pressure distributions indicated in FIG. 33 thatthe surface pressure P₂ acting on the upper internal leading wedge 671and upper trailing wedge 672 portions of the biplane is constant.Likewise, the pressure P₂ acting on the lower internal leading wedge 674and lower internal trailing wedge 676 is constant. In essence, thethrust vector component acting against the trailing wedge portions 672and 676 is equal to and cancels the drag vector component acting againstthe leading wedge portions 671 and 674. The supersonic wave or pressuredrag is created by and is equal to the pressure differential acting overan area, as a consequence of the even pressure distribution, there is nopressure or wave drag created by a properly designed biplane.

The biplane type structure 660 has been tested and proven for supersonicdrag reduction in other applications, and is ideally suited for use inconstruction of a rotor in the present invention.

In addition to the pressure drag which is advantageously eliminated bythe above described technique, there are unavoidable viscous drageffects created by the momentum lost in the boundary layers on theinternal and external surfaces of the biplane 660. However, such draglevels are small compared to the wave or pressure drag, and an exactanalysis including such effects may be conducted by those trained in theart.

Another advantage of the biplane type rotor is that the drag experiencedat off-design mach numbers gradually increases from that experienced atthe design point, i.e. there is no sudden jump in drag when the speedincreases or decreases somewhat from the design point. This isillustrated in FIG. 34, where the pressure diagram shows spikes ofoverpressure and of underpressure which correspond to the pressureacting against the trailing wedge portions 672 and 676 and to theleading wedge portions 671 and 674 at the corresponding chordwiselocation along the airflow path. Since supersonic wave or pressure dragis created by and is equal to the pressure differential acting over anarea, the consequence of an uneven pressure distribution is that thehigh pressure and low pressure spikes acting on small surface areaportions create small increases in drag.

Although FIG. 33 Illustrates the desired pressure profile through theinterior of the biplane 660, the specifics which are necessary to effectsuch a uniform spanwise pressure profile during rotor operation areshown in Table I for one embodiment of the invention, generally depictedin FIG. 35. Table I should also be viewed in conjunction with FIG. 6(which is illustrative of the basic layout and dimensions of a biplanerotor).

TABLE I Radial Biplane Section Gap Location Height Chord Thickness AreaHeight (in) (in) (in) (in) (in) (in) 36.000 5.66 12.000 1.911 22.9301.838 34.200 5.984 11.933 1.990 23.744 2.004 32.400 6.308 11.833 2.05424.410 2.200 30.898 6.578 11.856 2.095 24.834 2.389 28.296 7.046 11.8432.130 25.224 2.786 25.704 7.514 11.884 2.108 25.043 3.299 23.400 7.92811.977 2.023 24.224 3.883 22.104 8.162 12.059 1.940 23.394 4.282 20.6028.432 12.186 1.806 22.008 4.820 18.000 8.800 12.080 1.738 21.000 5.32315.429 9.200 12.210 1.750 21.368 5.700 12.857 9.500 12.400 1.800 22.3205.900 10.286 9.750 12.600 1.850 23.310 6.050 7.714 9.990 12.800 1.90024.320 6.190 5.143 10.100 13.100 1.980 25.938 6.140 2.571 10.210 13.4002.050 27.470 6.110 0.000 10.210 13.500 2.200 29.700 5.810

It can be seen from an examination of Table I that the gap height “G”,the distance between the upper 668 u and lower 668(l) shoulders of therotor 660, varies along the span of the rotor. This spanwise variationis required to maintain optimal biplane performance (shock cancellation)over as much of the rotor as possible. The thickness “T” is the distancebetween the outer surface (lower 666 or upper 664) and the lower 668(l)or upper 668(u) shoulder, respectively. As with gap height “G”, biplanethickness “T” and chord vary spanwise so as to assist in establishingthe desired pressure profile within the biplane. Biplane height, thedistance between the upper surface 664 and the lower surface 666, variesalong the span of the rotor as indicated in FIG. 33.

Biplane geometry supports optimal aerodynamic performance over the highspeed region of the outer rotor, but is not as effective on the inboardrotor sections, due to the decreasing Mach number.

The spanwise variation in rotor geometry reflected in Table I results inoptimal biplane performance (constant internal pressure) over theoutboard 12-14 inches of that rotor, and a desireable total geometricenvelope to house structural load bearing material on the inboard 22-24inches.

The relationships between rotor section geometry as defined by chord,gap height and thickness which result in the shock structure required toachieve shock cancellation (as described in connection with FIG. 33above) are well developed in a number of texts treating supersonic fluiddynamics and are, in general, well known or easily determined by thoseskilled in the art.

In summary, Table I sets forth an exemplary spanwise variation in rotorgeometry including chord, thickness and air gap which result indesirable rotor performance. Overall rotor size is not a criticalparameter and can easily be varied within a range of reasonable valuesas long as specific section geometry is maintained so as to result innecessary interior shock geometry and cancellation.

Alternate Rotor Shapes

Although the just described biplane rotor means offers significantadvantages for connecting a thrust module to a rotating shaft, otherrotor shapes may also substantially accomplish the desired results,namely, provide adequate strength with a minimum of drag.

Reference is now made to FIG. 41, where one viable alternativeconfiguration for a rotor 680 is illustrated. A continuous disc 682,with the thrust modules 684 mounted on the rim 686, is provided. It isimmediately clear that there can be no flow through a given section asin the case of the discrete biplane type rotor discussed above. However,a simple flat disc rotating in an otherwise quiescent fluid does havedrag. More accurately, the viscous interaction between the rotating discand the fluid surrounding the disc consumes power.

Returning now to FIG. 36, a schematic representation is provided of theflow field which results when a disc 700 is caused to rotate in anotherwise motionless fluid. As indicated in FIG. 36, the layer of airnear the disc 700 is carried by the disc 700 in the direction of arrowsreferenced 702 through friction and such air is thrown outward in thedirection of arrows 704 owing to the action of centrifugal forces. Thisradial airflow results in replacement airflow downward in an axialdirection (along the “z” axis) towards the disc 700. Once thereplacement air enters the flowfield, it is then carried downward towardthe disc 700 and is ejected centrifugally in the direction of arrows704, like the air which it is replacing. Thus it is seen that in thecase of a disc 700 rotating in air, there exists airflow in threedimensions: (a) in the radial direction, r, (b) in the circumferentialdirection, w (indicated by reference numeral 702), and (c) in the axialdirection, z.

The momentΓrequired to turn such a disc 700, wetted on both sides,(i.e., the force required to overcome drag on the disc turning at thedesired speed about the fixed axis z) has been shown to be$\begin{matrix}{{2\Gamma} = {0.616\quad {\pi\rho}\quad R^{4}\sqrt{{vw}^{3}}}} & (23)\end{matrix}$

FIG. 37 shows a plot of the rotating disc moment coefficient versusReynolds number, where the dimensionless moment coefficient is definedas $\begin{matrix}{C_{m} = \frac{2\Gamma}{\frac{1}{2}\rho \quad w^{2}R^{5}}} & (24)\end{matrix}$

and where the Reynolds number “” is defined as follows: $\begin{matrix}{ = \frac{R^{2}w}{v}} & (25)\end{matrix}$

This plot assumes the disc to be rotating in free air. Curve A in FIG.37 is the behavior predicted for laminar flow over the disc, asdescribed by equation (25) below: $\begin{matrix}{C_{m} = \frac{3.87}{\sqrt{}}} & (26)\end{matrix}$

The moment coefficient predictions for the transition region during achange in flow from laminar to turbulent is set forth in curve B, and isbased on empirical data. For turbulent flow at higher Reynolds numbers“”, the moment coefficient is better described by curves C and D. CurveC is described as follows: $\begin{matrix}{{- \frac{1}{\sqrt{C_{m}}}} = {{1.97\quad {\log \left( {\sqrt{C_{m}}} \right)}} + 0.03}} & (27)\end{matrix}$

Curve D is described as follows: $\begin{matrix}{C_{m} = {0.146^{- \frac{1}{5}}}} & (28)\end{matrix}$

The difference between curves C and D results from slightly differentassumptions with regard to behavior of the boundary layer profile, withcurve C assuming a logarithmic profile, and with curve D assuming a{fraction (1/7)}th power law profile.

One alternative which has also been considered is the inclusion of ahousing around the disc so as to reduce the moment coefficients and as aresult reduce undesirable parasitic power losses. The basic conceptillustrative of such a configuration is shown in FIGS. 38 and 39, wherea “tight” housing, i.e., one in which the gap “s” is small compared withthe radius “R” of the disk 802.

In the case of all laminar flow over the plate, the moment contributedby both sides of the disk can be described as follows: $\begin{matrix}{{2\Gamma} = {\pi \quad {wR}^{4}\frac{\mu}{s}}} & (29)\end{matrix}$

and the moment coefficient becomes $\begin{matrix}{C_{m} = {2\pi \frac{R}{s}\frac{1}{}}} & (30)\end{matrix}$

This equation is seen plotted as curve E in FIG. 40 for a ratio of gap“s” to radius “R” (s/R) of the disk 802 of 0.02. As the Reynolds numberincreases, the flow begins to transition from laminar to turbulent asindicated by the following approximate relation which appears as curve Gon FIG. 40. $\begin{matrix}{C_{m} = {2.67\quad ^{- \frac{1}{2}}}} & (31)\end{matrix}$

With increasing Reynolds numbers, the flow becomes fully turbulent andthe moment coefficient is more accurately represented by the followingequation: $\begin{matrix}{C_{m} = {0.0622\quad ^{- \frac{1}{5}}}} & (32)\end{matrix}$

which appears as curve H in FIG. 40. It can be seen that the provisionof a housing provides an order of magnitude or more reduction incoefficient of moment, by simple comparison of curve D of FIG. 37 withcurve H of FIG. 40.

Since it is not practical to completely enclose the thrust modules, apartial housing can provide a significant measure of the benefits of ahousing, and may be utilized as appropriate. In FIG. 42, use of such apartial upper housing 804 and lower housing 806 is shown used inconjunction with the flat, solid disc 682 illustrated in FIG. 41.

Another alternate embodiment is shown in FIGS. 43 and 44, where partialupper 810 and lower 812 housings closely hug the sloping or taperedsolid disc rotor 814. Thrust modules 816 are attached to the rim 818 ofrotor 814, to turn rotor 814 and shaft portions 820 and 822.

Yet another alternate embodiment is shown in FIGS. 45 and 46, where acombination of the prior concepts is utilized. Here, a small solid discrotor portion 830 has affixed thereto two or more biplane type arms 832,each of which has an upper 834 and lower 835 portion similar to thebiplane rotors described above. The disc 830 may be either flat ortapered as set forth in either FIG. 41 or FIG. 43. Thrust modules 836are affixed to the distal ends 338 of biplane arms 832, to drive theentire assembly so as to rotate shaft portions 840 and 842. To reducedrag, partial upper 844 and lower 846 housings are provided.

Rotor Materials of Construction

The structural design and material systems used for the rotor means arecritical to this power plant system. The rotor structure is as importantas the aerodynamic performance of the rotor and the propulsiveperformance of the thrust module discussed above. All three designelements (rotor materials, rotor aerodynamic design,and thrust moduleperformance) must be properly executed to place into operation a highperformance, maximum efficiency power plant as set forth herein.

Because of the centrifugal loads induced by the extreme speed with whichthe rotor turns, the material and structural characteristics of therotor are vitally important design elements. The following backgrounddiscussion is offered to illustrate the magnitude of the forcesinvolved.

Consider two basic elements rotating about an axis. FIG. 47 shows anuntapered rod 850 rotating about an axis perpendicular to its ownlongitudinal axis. FIG. 48 shows an untapered disc 852 rotating about anaxis through its center. In the case of the rod 850, the maximum stressoccurs at the center of the rod. Equation (33) below indicates thevariation in this peak stress with rotation rate: $\begin{matrix}{\sigma_{\max} = {\left( \frac{1}{2} \right)\frac{W}{386.4}\frac{{Lw}^{2}}{S}}} & (33)\end{matrix}$

In the case of the rotating disc 852, the maximum tangential and radialinertial stresses both occur at the center of the disc. The variation inthe magnitude of this stress with disc rotation rate “w” is indicated byequation (34): $\begin{matrix}{\sigma_{r} = {\sigma_{t} = {\sigma_{\max} = {\left( \frac{1}{8} \right)\frac{\delta \quad w^{2}}{386.4}\left( {3 + v_{p}} \right)R^{2}}}}} & (34)\end{matrix}$

These basic shapes are highly representative of two configurations forthe rotor means necessary for my power plant. The rod is analogous to anuntapered rotor arm, while the disk is obviously equivalent to theuntapered solid disc rotor configuration just discussed above. Bothequations (33) and (34) indicate stress levels expressed as pounds perinch (lbs/in), as a function of rotation rate.

It is instructive to consider the specific stress, that is, the stressper unit mass of material. The specific stress has units of inches,because the density of a specific material is cancelled out of themathematical relation. Thus, specific stress varies only with rotationrate. In FIG. 49, a plot of the variation of specific stress withrotation rate is shown for an untapered rod, and for a disc of uniformthickness (i.e., the shapes of FIGS. 47 and 48, respectively). It isimportant to note that at the rotation rates of importance in thepractice of the present invention, extremely high specific stresses areencountered, e.g., at a rotation rate of 15,000 rpm, about 1.5 millioninches of specific stress would be encountered by a rotating disc, andabout 1.8 million inches of specific stress would be encountered by arotating rod. It can be seen that in addition to the possibleaerodynamic advantages discussed above, a rotating disc also may offer aslight advantage with respect to materials requirements.

Any given material has associated with it a specific strength which iscommonly defined as the ultimate tensile strength of a material dividedby its density. Like specific stress, specific strength has the units ofinches. The two values are directly comparable; specific strength setsforth the load which a given material can withstand, and specific stresssets forth the load which a given material will encounter when in use ina given application.

Table II shows the specific strength for titanium, advanced metal matrixcomposites and carbon based conventional composites. Evaluation of themeaning of the specific strength data is straightforward. It is clearfrom FIG. 49 and Table II that as the rotational speed of either thesolid disc or the rod is increased, the specific stresses required mayultimately reach the specific strength of a given material. If the speedis increased beyond that point, the load will exceed the specificstrength, and as a result, the material will fail. In summary, thespecific stress expected to be encountered by rotors for the instantinvention exceeds the specific strength of commonly available materialssuch as steel, magnesium, and aluminum, and thus such materials are notsuitable for use as the primary structural material in the rotor meansof the present invention.

TABLE II SPECIFIC STRENGTHS FOR VARIOUS MATERIALS Material SpecificStrength (inches) Steel   176,000 Magnesium   584,610 Aluminum   594,060Titanium   683,220 Silicon Carbide Reinforced Titanium  1,300,250¹Kevlar Reinforced Polyester  3,752,600 Monofilament Carbon fibers15,000,000 ¹Silicon carbide casing on carbon fiber

The rotor means for the proposed power plant must turn at speeds between10,000 and 20,000 rpm. It is readily apparent from FIG. 49 and Table IIthat not even titanium, with its excellent specific strengthcharacteristics would represent a practical material for rotorconstruction. It is possible to reduce the specific stress by tapering agiven rotor element, and in fact, that was the approach used for analternate biplane rotor structure introduced herein below.

It is clear from relations (33) and (34), and can be easily visualizedfrom FIG. 49 and Table II, given the specific stress levels encounteredby rotor shapes operating at the speeds required, that commonly utilizedmetals or metal alloys do not have sufficient specific strength towithstand the loads encountered at the most desirable rotation rates.Newly developed metal matrix composites do provide acceptable strength,however, and can survive the required loads.

Carbon fiber reinforced polyester and epoxy composites easily have thespecific strength required for service in the instant invention. Asindicated in Table II, pure carbon monofilament fiber bundles or “tows”are commercially available with specific strength levels up to 15million inches. This is off the scale used for FIG. 49 and clearly has awealth of extra strength capability.

Unfortunately, when unprotected, both carbon fiber and epoxy compositeslack high temperature capability. However, if insulated from anoxidizing environment, the carbon tows can accommodate extremely hightemperature with only minimal reduction in strength.

In one embodiment, the basic rotor structure can be designed andfabricated using both metal matrix composites and carbon or other highstrength fiber windings. With proper thermal and oxidative protection,monofilament carbon fiber tows can be combined into a structure withexcellent strength and high temperature capability. In the compositedesign, high strength is provided by continuous monofilament carbonfibers, so as to give the structure sufficient reinforcement towithstand the centrifugal loads encountered. In that design, the metalmatrix composite shells or “gutters” provide the shape integrity andrigidity required for proper aerodynamic performance.

Because the carbon windings are far stiffer than the silicon-carbidereinforced titanium, the metal matrix composite elements as well as thethrust modules at the end of the rotors are restrained from excessiveand damaging deflection by the ultra-high strength carbon fiberwindings. The high specific strengths of the carbon fibers make themquite suitable for the fabrication of stiff, strong, and lightweightcomposite rotors which can minimize vibrational and static load bending.The carbon fiber windings thus become a central tensile reinforcementelement which carries the bulk of all centrifugally induced mechanicalloads.

FIG. 50 shows the spanwise variation in total rotor materialcross-sectional area for one exemplary embodiment of the biplane rotor,such as is set forth in FIGS. 5 and 6 above. Implicit in thecross-sectional area numbers provided in FIG. 50 is that the rotor istapered. The rotor taper ratio is defined as follows $\begin{matrix}{\lambda = \frac{A_{root}}{A_{tip}}} & (35)\end{matrix}$

The ultimate specific strength for the silicon-carbide reinforced metalmatrix composite and the ultimate specific strength of the carbonmonofilament is indicated in TABLE II. The spanwise variation in stressencountered by the silicon carbide reinforced metal matrix compositegutter portion in one embodiment of the rotor means (as set forth inFIGS. 5 and 6 above) of the present invention is set forth in FIG. 51.Likewise, the stress encountered by the carbon reinforcing fibers in thesame embodiment of the rotor means of the present invention is set forthin FIG. 52. While merely illustrative of the situation in one embodimentof the invention, the curves of FIGS. 51 and 52 assume a constantcumulative carbon fiber cross-sectional area of approximately 9 squareinches. Note the substantial safety margin in both cases.

The safety margin just identified can be further increased by increasingthe material taper ratio. Preferably, in order to minimize the actualloading to the extent practical, the rotor means should be built withhigh strength materials in shapes which have large material taperratios. This basically means that at increasing radial station, (furtherfrom the axis of rotation), the rotor means should become increasinglyslender or thin. Fundamentally, reduction of rotating mass results inreduction of the encountered stress operating at the center of rotation.

Attention is now directed to FIGS. 5 and 6, where reference numerals 224and 226 are the lower and upper biplanes. The gutter, metal matrixcomposite portion is denoted 860, and contains on either side of acentral rib 248 a, a pair of troughs 864, generally triangular in shape,adapted to receive the high strength carbon fibers 242 therein. Therotor gutter 860 is preferably provided in an advanced metal matrixcomposite, silicon-carbide reinforced titanium. Protection for thecarbon fibers 242 is provided by cover plate 260 and end caps 270(l) and270(t). Not shown in FIG. 5 is the symmetric mirror reflection of theother half of the rotor structure.

The thrust module 230 flowpath shape previously introduced (see FIG. 8and accompanying discussion above) was specifically developed toaccommodate the carbon fiber 242 windings. FIG. 8 shows how the carbonfibers are integrated into the radially outermost region of the thrustmodule. Note that the groove 870 into which the fibers are wound iswider than it is deep. This spreads the loads over as large an area aspossible and thus minimizes stress concentrations. Like the rotors, thethrust modules may be supplied with a central rib 872 for addedstiffness.

To assemble the biplane rotor 220 structure, the thrust modules 230 areplaced in position on the ends of the gutter 860 (i.e., the rotorsuperstructures) and then a continuous tow of carbon fiber filaments 242are progressively wound around the entire biplane/thrust moduleassembly. The carbon fiber tow 242 is held in constant tension as it issuccessively wound into the accommodating trough 864 in the gutter 860and similar groves 870 in the thrust modules 230. The tension in thecarbon fiber 242 acts to hold the entire assembly together. Furthermore,it pre-loads the rotor structure with an initial compressive load whichhelps to reduce operational stresses in the rotor. The beginning andending of the carbon fiber windings as well as any interim splices aresecured with epoxy, thermosetting or other suitable resin within therotor. After the winding is completed, the rotor top and bottom plates(top plate 260(l) and 260(t) has a counterpart on the bottom, 262(l) and262(t)), as well as the thrust module 230 end caps (part numbers 270(l)and 270(t) are bonded in place.

Alternate Rotor Design

The key to the current rotor structure lies in the use of high strengthmaterials. As an alternative to use of carbon fiber or other highstrength windings, a solid rotor design may be completed utilizingsilicon carbide coated carbon fiber metal matrix composite materials.Such a design is set forth in FIG. 55 (and related cross sections FIGS.A through G), and FIGS. 56 and 57. In FIG. 55 a partial isometric viewof a second embodiment of the biplane rotor of the present invention,similar to the view first set forth above in FIG. 5, is provided showinga solid metal matrix composite type construction configuration. Rotor920 is provided having thrust module 930 at the distal end thereof. Anupper solid rotor 932 and a lower solid rotor 934 are provided. Endcaps935 are used to secure thrust modules 930. The exact shape and size ofrotors 932 and 934 are determined by the same aerodynamic (uniformpressure profile) and strength (minimize stress to the extent possible)objectives as discussed in detail above. The upper 932 and lower 934rotors are secured in a central hub 936 having appropriate fasteners938. The interior of rotors 932 and 934 may be layers 940 of siliconcarbide reinforced titanium.

Fuel may be supplied to the thrust module 930 through conduit means942(u) (upper) and 942(l) (lower) which is defined by edge 944 in shaftmeans 946 and in rotors 932 and 934. The progress of fuel conduits942(u) and 944(l) spanwise through rotors 932 and 934 is depicted inFIGS. A through G.

FIG. A is a vertical cross-sectional view taken through line A—A of FIG.55, showing the construction of the solid type rotor.

FIG. B is a vertical cross-sectional view taken through line B—B of FIG.55, showing the construction of the solid type rotor, and also showingthe changing features of gap G and fuel conduit 942(u) and 942(l)diameter.

FIG. C is a vertical cross-sectional view taken through line C—C of FIG.55, similar to the view set forth in FIGS. A & B above, showing furthervariations in rotor dimensions with change in radial position.

FIG. D is a vertical cross-sectional view taken through line D—D of FIG.55, similar to the views in FIGS. A through C above, showing furthervariations in rotor dimensions with change in radial position.

FIG. E is a vertical cross-sectional view taken through line E—E of FIG.55, similar to the view set forth in FIGS. A through D above, showingfurther variations in rotor dimensions with change in radial position.

FIG. F is a vertical cross-sectional view taken through line F—F of FIG.55, similar to the views in FIGS. A through E above, showing furthervariations in rotor dimensions with change in radial position.

FIG. G is a vertical cross-sectional view taken through line G—G of FIG.55, similar to the views in FIGS. A through F above, showing furthervariations in rotor dimensions with change in radial position.

FIG. 56 provides an isometric view of an end cap 935 for use with thesolid type rotor 920 first illustrated in FIG. 55 above. In order tokeep the thrust module 930 in place while at the high centrifugal speedsof operation, cap 935 will be fused, brazed, welded, or otherwisemetallurgically bonded in a high strength joint via use of a series ofinterlocking tongues 950 and grooves 952. Groves 952 in cap 935accommodate complementary tongues 954 which are fashioned to the distalends of rotors 932 and 934, as more fully seen in FIG. 57. A series ofsteps K, L, and M may be provided in cap 935, complementary to receivingledges K′, L′, and M′ on the exterior of thrust module 930, so as toprovide the maximum possible tongue and groove surface to surfacecontact consistent with the necessary thrust module 930 internaldimensions.

FIG. 57 illustrates a vertical cross-sectional view of the finished,operating position of the end cap 935 just illustrated in FIG. 56, whenthe cap 935 is affixed to the upper 932 and lower 934 rotors.

Cogeneration System Configuration

Cogeneration refers to the simultaneous generation of electrical andthermal energy in a single powerplant. My powerplant can easilyincorporate both thermal and kinetic energy recovery withoutmodification to the basic configuration. To accommodate cogeneration,the horizontal annular exhaust duct would be configured as illustratedin FIGS. 1, 2, and 3, wherein the exhaust duct contained an integralheat exchanger. Coolant passages inside the duct would cool the hightemperature exhaust gases from the thrust module and heat the coolantflowing through the heat exchanger. If water was used as a coolant,coolant flowrate could be adjusted so as to generate high pressure steamfrom a continuous supply of water. This steam could be used as a sourceof heat or to drive a secondary steam turbine which in turn coulddirectly drive an electric generator.

Turbopower could be incorporated through the addition of a secondaryannular reaction gas turbine at the perimeter of the primary exhaust gasduct. The exhaust leaving the thrust module has both thermal and kineticenergy. The cogeneration system just described above would only capturethe thermal energy in the exhaust flow. However, a reaction turbinecould extract a large portion of the total available kinetic energy fromthe exhaust. The reaction turbine could then be used to providemechanical energy for other uses, or could drive a secondary electricgenerator. A reaction turbine could be placed either before or after aheat exchange section, but it is preferred to locate the turbine afterthe heat exchanger to extend service life of the materials in theexhaust gas stream.

Addition of Turbine to Exhaust

Attention is now directed to FIG. 53, where a vertical cross-sectionalview of the power plant 1000 of the present invention is provided,similar to the view first set forth in FIG. 1 above, but here showingthe addition of an annular reaction turbine 1002 for capturing thekinetic energy of the exhaust gases and generating shaft or electricalpower therefrom.

FIG. 54 is a cross-sectional view, taken across line 54—54 of FIG. 53,here showing the general position and airfoil shaped vanes 1004 of theannular reaction turbine 1002. The power plant 1000 as depicted in thisFIG. 54 shows a stationary annular exhaust duct 1006. Heat exchangeelements 1008 are located within this duct to remove heat energy fromthe exhaust gases and to transfer that beat to a secondary working fluidsuch as water. The water can be heated to high pressure steam, and canthereafter be used: (a) to drive a steam turbine, for (i) shaft work or(ii) to drive an electrical generator, or (b) as process heat. However,even under ideal conditions, the heat exchanger can only remove thethermal energy from the exhaust gas stream.

Due to its high velocity, there is a large amount of kinetic energy inthe exhaust gas stream 1001. An annular reaction turbine rotor 1002mounted around the outer circumference of the exhaust duct 1003 extractsa portion of the kinetic energy remaining in the flow when it exits thefixed exhaust from the heat exchange section. The airfoil shaped turbinevanes 1004 utilize the high velocity exhaust gas flow 1001 to generate aforce which drives the reaction turbine 1002 in the direction indicatedby reference arrow 1010. The motion of the turbine can then be convertedto mechanical or electrical energy by way of ring gear 1012 mounted onthe exterior of the reaction turbine 1002. The ring gear 1002 isconnected to a driven gear 1014, gearbox 1016, and generator 1018 toaccomplish the desired electrical generation.

Including a reaction turbine 1002 makes it possible to extract thekinetic energy remaining in the exhaust gas stream and convert it tousable energy. This further enhances the performance characteristics ofmy power generating plant design.

FIG. 58 shows schematically the use of the power and heat generated inthe thrust module of the power plant for a variety of heat recovery,shaft work, or electrical generation activities. These options are asjust described, with respect to use of the reaction turbine. Otheroptions have been set forth above.

Both FIGS. 1 and 2 above show my powerplant design with a cogenerationsteam extraction system. This system includes a stationary exhaust ductwhich surrounds the rotor, with the duct filled with hollow vanesthrough which a secondary working fluid is circulated. In the currentdesign the working fluid to be circulated through the duct is water,although it is clear that a variety of heating fluids could be utilizedby those skilled in the art. As shown, the hot exhaust gases from thethrust module heat the duct and the water, ultimately generating highpressure steam which could be used to drive a secondary steam turbine orturbines. However, the instant power plant is sufficiently efficientthat it can be operated cost effectively without the exhaust enthalpyextraction system.

Power Plant Efficiency and Performance

In FIG. 59, the performance of my ACRE™ power plant is compared to theperformance of a gas turbine power plant, in terms of heat rate (“HR”).Variations in performance of the instant invention are shown as afunction of the characteristic speed or ramjet Mach number. The heatrate is a performance term typically used by powerplant designers. TheHR is the amount of heat added, usually in BTU units, required toproduce a unit of work output, usually expressed in kilowatt hours (kwh)or horsepower-hours (Hp-hr). Heat rate is inversely proportional to thesystem efficiency, hence the lower the HR value, the better. Since fuelcost is typically known in terms of BTU heating value and the value ofelectrical power generated is capable of being forecast for variousplant locations and situations, the heat rate allows the calculation ofthe economic viability of a given system.

Conventional gas turbine systems produce power in the range of roughly8,000 to 11,000 BTU/Hp-hr. The instant invention, without co-generationor reaction turbine, is projected to produce power in the range ofslightly less than 5,000 to about 7,000 BTU/Hp-hr, with an optimumcurrent design of between 5,500 and 5,700 BTU/Hp-hr. Simply stated,conventional gas turbines require from about 40% to about 100% more fuelto produce the same amount of electricity than the instant invention.

The power plant performance characteristics: (a) with cogeneration, (b)without cogeneration, and (c) with use of a reaction turbine, are shownin FIG. 60. FIG. 60 assumes a tip Mach number of 3.5 and represents thevariation in performance with changing throttle settings. From FIG. 60it is clear that the plant achieves optimal performance at a lowthrottle setting. Moreover, cogeneration improves the basic plantperformance by approximately 32%, resulting in heat rates of about 4,000BTU/Hp-hr. Significantly, the addition of an annular reaction turbineadds another 28% to the efficiency rating, to allow heat rates in the2,500 BTU/Hp-hr range.

From FIG. 59 it is clear that my power plant significantly out-performsconventional gas turbine systems. The gas turbine industries are quitemature, and manufacturers have been refining and improving turbinesystems for about half a century. In general, contemporary increases ingas turbine performance are very small; most increases are measured infractions of a percent efficiency. Thus, an overall output available atthe high efficiency, low heat rate levels indicted on FIG. 59 show thatmy power generation apparatus and method provides a major, fundamentalimprovement in overall power generation economics.

Another way of expressing the efficiency of the instant invention isshown in FIG. 61. Thermal efficiencies (the ratio of fuel energy inputto the mechanical energy output) for various types of power plants isillustrated. For illustrative purposes, it can be said that piston typeengines fall in the 30% efficiency range, gas turbines in the 40%efficiency range, and the baseline ramjet power plant falls in the 50%efficiency range. Use of co-generation and an annular reaction turbinefurther improves the cycle efficiencies of the instant invention up tothe 70% range. The advantages of the instant invention are thusself-evident.

The cost of new electrical generation capacity using my power plant andcomparisons with other types of power plants, is shown in FIG. 62. Ascurrently understood, it is expected that the cost of electricityproduced by a basic ramjet driven power plant, as described herein,including both capital and operating costs, will be in She range of$0.02 per Kwh, which is some 50% or more less than the cost of powerfrom currently known power plants.

The method and apparatus for producing mechanical, electrical, andthermal power as described above provides a revolutionary, compact,easily constructed, cost effective power plant. The output from thispower plant can be used in conjunction with existing power deliverysystems, and represents a significant option for reducing air emissionsby combustion of clean burning fuels. Further, given the efficiencies,dramatically less fuel will be consumed per unit of electrical,mechanical, or thermal energy generated.

It will thus be seen that the objects set forth above, including thosemade apparent from the proceeding description, are efficiently attained,and, since certain changes may be made in carrying out the above methodand in construction of the apparatus and in practicing the methods setforth without departing from the scope of the invention, it is to beunderstood that the invention may be embodied in other specific formswithout departing from the spirit or essential characteristics thereof.The present embodiments are therefore to be considered in all respectsas illustrative and not as restrictive. Accordingly, the scope of theinvention should be determined not by the foregoing description and theembodiments illustrated, but by the appended claims, and consequentlyall changes, variations, and alternative embodiments which come withinthe meaning and range of equivalents of the appended claims aretherefore intended to be embraced therein.

Appendix 1

List of Equations $\begin{matrix}{M = \frac{v}{a}} & (1.1) \\{F = {{\left( {{\overset{.}{m}}_{0} - {\overset{.}{m}}_{1}} \right)v_{7}} + {\left( {{\overset{.}{m}}_{1} + {\overset{.}{m}}_{f}} \right)v_{6}} - {{\overset{.}{m}}_{0}v_{0}}}} & (1.2) \\{F = {{{\overset{.}{m}}_{1}\left( {v_{6} - v_{0}} \right)} + {{\overset{.}{m}}_{f}v_{6}} - {\left( {{\overset{.}{m}}_{0} - {\overset{.}{m}}_{1}} \right)\left( {v_{0} - v_{7}} \right)}}} & (1.3) \\{A_{0} = A_{1}} & (1.4) \\{A_{5} = A_{6}} & (1.5) \\{F = {{{\overset{.}{m}}_{1}\left( {v_{6} - v_{0}} \right)} + {{\overset{.}{m}}_{f}v_{6}}}} & (1.6) \\{F = {{{\overset{.}{m}}_{1}\left( {v_{6} - v_{1}} \right)} + {{\overset{.}{m}}_{f}v_{6}} + \left( {{p_{6}A_{6}} - {p_{1}A_{1}}} \right)}} & (1.7) \\{T = {{{pA} + {\overset{.}{m}v}} = {{pA}\left( {1 + {\gamma \quad M^{2}}} \right)}}} & (1.8) \\{T = {{\overset{.}{m}\left( \sqrt{\frac{2\left( {\gamma + 1} \right)}{\gamma}R_{g}T_{t}} \right)}{\varphi (M)}}} & (1.9) \\{{\varphi (M)} = \frac{1 + {\gamma \quad M^{2}}}{\sqrt{2\left( {\gamma + 1} \right){M^{2}\left( {1 + {\frac{\left( {\gamma - 1} \right)}{2}M^{2}}} \right)}}}} & (1.10) \\{{{\overset{.}{w}}_{f}S_{f}} = {{{\overset{.}{w}}_{a}S_{a}} = {\overset{.}{m}\sqrt{\frac{2\left( {\gamma + 1} \right)}{\gamma}R_{g}T_{t}}}}} & (1.11) \\{p_{t5} = {p_{t0} \times \frac{p_{t1}}{p_{t0}} \times \frac{p_{t2}}{p_{t1}} \times \frac{p_{t4}}{p_{t2}} \times \frac{p_{t5}}{p_{t4}}}} & (1.12) \\{T_{5} = {p_{0}{A_{5}\left( {1 + {\gamma_{5}M_{5}^{2}}} \right)}}} & (1.13) \\{T_{5} = {{\overset{.}{w}}_{a}S_{a}{\varphi \left( M_{5} \right)}}} & (1.14) \\{T_{1} = {p_{0}{A_{1}\left( {1 + {\gamma_{0}M_{0}^{2}}} \right)}}} & (1.15) \\{{f\left( M_{0} \right)} = \frac{{\overset{.}{w}}_{a}\sqrt{T_{t0}}}{p_{0}A_{1}}} & (1.16) \\{T_{1} = {\frac{{\overset{.}{w}}_{a}\sqrt{T_{t0}}}{f\left( M_{0} \right.}\left( {1 + {\gamma_{0}M_{0}^{2}}} \right)}} & (1.17) \\{{T_{5} - T_{1}} = {{{\overset{.}{w}}_{a}S_{a}{\varphi \left( M_{5} \right)}} - {\frac{{\overset{.}{w}}_{a}\sqrt{T_{t0}}}{f\left( M_{0} \right)}\left( {1 + {\gamma_{0}M_{0}^{2}}} \right)}}} & (1.18) \\{D = {\left( {p_{2} - p_{3}} \right)t}} & (1.19) \\{L = {\left( {p_{2}^{\prime} - p_{2}} \right)c\quad \cos \quad \beta}} & (1.20) \\{D = {\left( {p_{2}^{\prime} - p_{2}} \right)c\quad \sin \quad \beta}} & (1.21) \\{P_{ower} = {\int_{hub}^{tip}{r\quad {wD}{r}}}} & (1.22) \\{{2\Gamma} = {0.616\quad {\pi\rho}\quad R^{4}\sqrt{{vw}^{3}}}} & (1.23) \\{C_{m} = \frac{2\Gamma}{\frac{1}{2}\rho \quad w^{2}R^{5}}} & (1.24) \\{\Re = \frac{R^{2}w}{v}} & (1.25) \\{C_{m} = \frac{3.87}{\sqrt{\Re}}} & (1.26) \\{{- \frac{1}{\sqrt{C_{m}}}} = {{1.97\quad {\log \left( {\Re \sqrt{C_{m}}} \right)}} + 0.03}} & (1.27) \\{C_{m} = {0.146\quad \Re^{- \frac{1}{5}}}} & (1.28) \\{{2\Gamma} = {\pi \quad {wR}^{4}\frac{\mu}{s}}} & (1.29) \\{C_{m} = {2\pi \frac{R}{s}\frac{1}{\Re}}} & (1.30) \\{C_{m} = {2.67\quad \Re^{- \frac{1}{2}}}} & (1.31) \\{C_{m} = {0.0622\quad \Re^{- \frac{1}{5}}}} & (1.32) \\{\sigma_{\max} = {\left( \frac{1}{2} \right)\frac{W}{386.4}\frac{{Lw}^{2}}{S}}} & (1.33) \\{\sigma_{r} = {\sigma_{t} = {\sigma_{\max} = {\left( \frac{1}{8} \right)\frac{\delta \quad w^{2}}{386.4}\left( {3 + v_{p}} \right)R^{2}}}}} & (1.34) \\{\lambda = \frac{A_{root}}{A_{tip}}} & (1.35)\end{matrix}$

Appendix 2

Nomenclature

F=net thrust module thrust

T=stream thrust

T₅=stream thrust at thrust module station (eg., see FIG. 25)

T₁=stream thrust at thrust module station (eg., see FIG. 25)

{dot over (m)}₀=mass flow through captured stream tube

{dot over (m)}₁=mass flow through thrust module station 1 (eg., see FIG.25)

{dot over (m)}_(f)=mass flow of fuel into thrust module (eg., see FIG.25)

{dot over (w)}_(a)=weight flow of air into thrust module ({dot over(m)}₀g)

{dot over (w)}_(f)=weight flow of fuel into thrust module ({dot over(m)}_(f)g)

v=velocity

v₀=velocity of fluid at thrust module station 0 (eg. see FIG. 25)

v₆=velocity of fluid at thrust module station 6 (eg., see FIG. 25)

v₇=velocity of fluid at thrust module station 7 (eg., see FIG. 25)

A₀=cross sectional flow area at thrust module station 0 (eg., see FIG.25)

A₁=cross sectional flow area at thrust module station 1 (eg., see FIG.25)

A₅=cross sectional flow area at thrust module station 5 (eg., see FIG.25)

A₆=cross sectional flow area at thrust module station 6 (eg., see FIG.25)

p=static pressure

p₀=static pressure at thrust module station 0 (eg., see FIG. 25)

p₁=free stream static pressure (eg., see FIGS. 28, 29 and 30)

p₂=post shock leading edge static pressure (eg., see FIGS. 28, 29 and30)

γ=ratio of specific heats

γ₀=ratio of specific heats at station 0 (eg., see FIG. 25)

γ₅=ratio of specific heats at station 5 (eg., see FIG. 25)

M=Mach number

M₅=Mach number at thrust module station 5 (eg., see FIG. 25)

T_(t)=stagnation temperature

T_(t0)=stagnation temperature at flow station 0 (eg., see FIG. 25)

R_(g)=ideal gas constant

φ(M)=Mach number function

S_(f)=fuel specific impulse

S_(a)=air specific impulse

p_(t)=total pressure

p_(t5)=total pressure at station 5 (eg., see FIG. 25)

p_(t0)=total pressure at station 0 (eg., see FIG. 25)

p_(t1)=total pressure at station 1 (eg., see FIG. 25)

p_(t2)=total pressure at station 2 (eg., see FIG. 25)

p_(t4)=total pressure at station 4 (eg., see FIG. 25)

p_(t5)=total pressure at station 5 (eg., see FIG. 25)

a=speed of sound

L=lift

D=drag

c=chord of airfoil section

α₀=angle of attack of airfoil section

β=angle of attack of flat plate (eg. see FIG. 30)

t=thickness of airfoil section (eg., see FIG. 28)

Γ=moment required to turn disc (eg., see FIGS. 36, 38 and 39)

F_(i)=radius of disc (eg., see FIG. 36, 38 and 39)

γ=local radial station

ν=specific viscosity of fluid surrounding the disc

μ=Kinematic viscosity of fluid surrounding the disc

ρ=density of fluid surrounding the disc

C_(m)=dimensionless moment coefficient

s=gap between housing and disc

ω=rotation rate

=Reynolds Number

σ_(max)=maximum stress

σ_(r) =radial stress

σ_(t)=tangential stress

S=cross-sectional area of rod

W=weight of rod

L=length of rod

ν_(p)=Poisons ratio

δ=material density

λ=material taper ratio

A_(root)=material cross-sectional area at rotor root

A_(tip)=material cross-sectional area at rotor tip

What is claimed is:
 1. An apparatus for generating power, comprising:(a) a support structure, said support structure comprising (i) anoxidant supply conduit, and (ii) a first housing portion with a rotorside surface, and (iii) a second housing portion with a rotor sidesurface; (b) a first output shaft, said first output shaft rotatablysecured with respect to said support structure; (c) a rotor having aradius R, wherein said rotor (i) is connected to said first output shaftto provide rotary motion of said first output shaft upon rotation ofsaid rotor, (ii) comprises at least one material having a specificstrength capability in excess of 683,220 inches, (iii) comprises a firstsurface portion, said first surface portion rotatably positioned in aclose fitting, first spaced apart relationship adjacent to said rotorside surface of said first housing portion, and (iv) comprises a secondsurface portion, said second surface portion rotatably positioned in aclose fitting, second spaced apart relationship adjacent to said rotorside surface of said second housing portion, and (v) wherein each ofsaid first and said second spaced apart relationships are defined by agap width “s” which is small compared to radius “R” of said rotor, to atleast a partially house said rotor in a tight fitting relationship, soas to minimize aerodynamic drag on said rotor; (d) one or more ramjetthrust modules, said one or more ramjet thrust modules (i) each securedto said rotor for rotation therewith, (ii) each further comprise aninlet and an outlet, and wherein said inlet and said outlet aresubstantially aligned in a linear configuration with respect to an inletairflow, (iii) each further comprise a shaped external portion, saidshaped external portion comprising a substantially constantcross-sectional size when sequentially examined in cross-sectionperpendicular to the axis of an inlet airflow from a forwardcross-section to a rearward cross-section, to thereby minimize pressuredrag when said one or more ramjet thrust modules operate at an inletairflow velocity M₀ of at least Mach 1.5, and (iv) each of which mixesfuel supplied thereto with an oxidant supplied via said oxidant supplyconduit in said support structure, to burn said fuel to generate hotcombustion gas which escapes from said one or more ramjet thrustmodules, producing thrust, thereby propelling said one or more ramjetthrust modules to turn said rotor and said first output shaft, thusproviding shaft power output from said apparatus.
 2. An apparatus forgenerating power, as set forth in claim 1, further comprising a heatrecovery section, said heat recovery section arranged to receive saidhot combustion gas from said one or more ramjet thrust modules, saidheat recovery section further comprising an inlet, an outlet, and asecondary working fluid for circulation to and from said heat recoverysection, whereby said hot combustion gas is cooled by recovering heattherefrom and transferring such recovered heat into said secondaryworking fluid.
 3. The apparatus as set forth in claim 2, wherein saidsecondary working fluid is used to provide thermal energy.
 4. Theapparatus as set forth in claim 2, wherein said secondary working fluidcomprises water, and wherein upon heating of said secondary workingfluid, steam is produced.
 5. The apparatus of claim 4, furthercomprising a steam turbine, wherein said steam which results fromheating of said water is contained under pressure and fed to said steamturbine to produce useful work on a steam turbine output shaft.
 6. Theapparatus as set forth in claim 5, wherein said steam turbine outputshaft is operatively connected to a first electrical generator, andwherein said useful work on said steam turbine output shaft turns saidfirst electrical generator to produce electricity.
 7. The apparatus asset forth in claim 5, further comprising a second electrical generator,and wherein said shaft work produced by said steam turbine output shaftturns said second electrical generator to produce electric power.
 8. Theapparatus as set forth in claims 6 or 7, wherein said apparatus consumesless than about 4,200 BTU/Hp-hr, based on combined cycle operation andthe ratio of fuel energy input to electrical energy output.
 9. Theapparatus as set forth in claims 6 or 7, wherein said apparatus consumesless than about 4,000 BTU/Hp-hr, based on combined cycle operation andthe ratio of fuel energy input to electrical energy output.
 10. Theapparatus for generating power as set forth in claim 2, wherein saidheat recovery section further comprises a plurality of aerodynamicallyshaped structures located substantially within said hot combustion gasflow path, said shaped structures adapted to allow said secondaryworking fluid to pass through the interior thereof in a heat exchangerelationship with said hot combustion gas which passes through said hotcombustion gas flow path in which said structures are located.
 11. Theapparatus as set forth in claim 1, wherein said first output shaft isoperatively connected to a first electrical generator, and wherein saidmechanical work provided at said first output shaft turns said firstelectrical generator to produce electricity.
 12. The apparatus as setforth in claim 1, wherein said apparatus generates shaft power at asimple cycle efficiency of at least 37 percent based on the ratio offuel energy input to mechanical energy output, when operating at aninlet velocity of at least Mach
 3. 13. The apparatus as set forth inclaim 1, wherein said apparatus generates shaft power at a simple cycleefficiency of at least about 45 percent, based on the ratio of fuelenergy input to mechanical energy output, when operating at an inletvelocity of at least Mach 3.5.
 14. The apparatus as set forth in claim1, wherein said apparatus generates shaft power at a simple cycleefficiency of at least 52 percent, based on the ratio of fuel energyinput to mechanical energy output, when operating at an inlet velocityof at least Mach
 4. 15. The apparatus as set forth in claim 1, whereinsaid apparatus consumes less than about 7,000 BTU/Hp-hr, based on simplecycle operation and the ratio of fuel energy input to mechanical energyoutput.
 16. The apparatus as set forth in claim 1, wherein saidapparatus consumes less than about 5,700 BTU/Hp-hr, based on simplecycle operation and the ratio of fuel energy input to mechanical energyoutput.
 17. The apparatus as set forth in claim 1, wherein saidapparatus operates at in inlet velocity of approximately Mach 3.5, andwherein said apparatus consumes between about 5,500 to about 5,700BTU/Hp-hr, based on simple cycle operation and the ratio of fuel energyinput to mechanical energy output.
 18. The apparatus of claim 1, whereinsaid rotor comprises a bi-plane.
 19. The apparatus of claim 18, whereinsaid bi-plane comprises opposing upper and lower bi-plane elements, andwherein each of said upper and lower bi-plane elements further comprisesa leading edge and a trailing edge with respect an airstream throughwhich said biplane is passing.
 20. The apparatus of claim 19, whereineach of said bi-plane elements are generally triangular in shape. 21.The apparatus of claim 20, wherein said bi-plane elements are situatedin a pre-selected profile, said pre-selected profile selected to providea substantially uniform pressure profile within said biplane, saiduniform pressure profile being maintained between said leading edge andsaid trailing edges, so as to cancel shock waves created by the movementof said bi-plane through said an airstream at supersonic speed, so as tominimize aerodynamic drag on said rotor.
 22. The apparatus of claim 18,wherein said oxidant supply conduit further comprises an annular shapedpassageway portion disposed for airflow movement at a rate sufficient sothat said bi-plane rotor is not significantly affected by turbulencefrom a wake created by said bi-plane rotor as it rotates through saidairflow in said oxidant supply conduit.
 23. The apparatus of claim 1,wherein said rotor comprises a central disc.
 24. The apparatus of claim1, wherein said rotor comprises (a) a central disc, and (b) a bi-plane,said bi-plane radially extending from said central disc.
 25. Anapparatus for generating power, comprising: (a) a support structure,said support structure comprising (i) an oxidant supply conduit, and(ii) a first housing portion with a rotor side surface, and (iii) asecond housing portion with a rotor side surface; (b) a first outputshaft, said first output shaft rotatably secured with respect to saidsupport structure; (c) a rotor, wherein said rotor (i) is connected tosaid first output shaft to provide rotary motion of said first outputshaft upon rotation of said rotor, (ii) comprises at least one materialhaving a specific strength capability in excess of 683,220 inches, (iii)comprises a first surface portion, said first surface portion rotatablypositioned in a close fitting, first spaced apart relationship adjacentto said rotor side surface of said first housing portion, and (iv)comprises a second surface portion, said second surface portionrotatably positioned in a close fitting, second spaced apartrelationship adjacent to said rotor side surface of said second housingportion, and (v) wherein each of said first and said second spaced apartrelationships are defined by a gap width “s” which is small compared toradius “R” of said rotor, to at least a partially house said rotor in atight fitting relationship, so as to minimize aerodynamic drag on saidrotor; (d) one or more ramjet thrust modules, said one or more ramjetthrust modules (i) each secured to said rotor for rotation therewith,(ii) each further comprising an inlet and an outlet, and wherein saidinlet and said outlet are substantially aligned in a linearconfiguration with respect to an inlet airflow, (iii) each furthercomprising a shaped external portion, said shaped external portioncomprising a substantially constant cross-sectional size whensequentially examined in cross-section perpendicular to the axis of aninlet airflow from a forward cross-section to a rearward cross-section,to thereby minimize pressure drag when said one or more ramjet thrustmodules operate at an inlet airflow velocity M₀ of at least Mach 1.5,and (iv) each of which mixes fuel supplied thereto with an oxidantsupplied via said oxidant supply passageway in said support structure,to burn said fuel to generate hot combustion gas which escapes from saidone or more ramjet thrust modules, producing thrust, thereby propellingsaid one or more ramjet thrust modules to turn said rotor and said firstoutput shaft, thus providing shaft power output from said apparatus; (e)a heat recovery section, said heat recovery section arranged to receivesaid hot combustion gas from said one or more ramjet thrust modules,said heat recovery section further comprising (i) a heat recovery inlet,(ii) a heat recovery outlet, and (iii) a secondary working fluid forcirculation to and from said heat recovery section, (iv) whereby saidhot combustion gas is cooled by recovering heat therefrom andtransferring such recovered heat into said secondary working fluid; (f)a steam turbine, wherein said secondary working fluid comprises water,and wherein said heating of said secondary working fluid produces steamthat is contained under pressure and fed to said steam turbine toproduce useful work on a steam turbine output shaft; (g) a firstelectrical generator, said first electrical generator operativelyconnected to said first output shaft, and wherein said mechanical workprovided at said first output shaft turns said first electricalgenerator to produce electricity; and (h) wherein said steam turbineoutput shaft is operatively connected to an electrical generator toproduce electricity.
 26. The apparatus as set forth in claim 25, whereinsaid apparatus generates shaft power from said first output shaft at asimple cycle efficiency of at least about 37 percent, based on the ratioof fuel energy input to mechanical energy output from said first outputshaft, when operating at an inlet velocity of at least Mach
 3. 27. Theapparatus as set forth in claim 25, wherein said apparatus generatesshaft power from said first output shaft at a simple cycle efficiency ofat least about 45 percent, based on the ratio of fuel energy input tomechanical energy output from said first output shaft, when operating atan inlet velocity of at least Mach 3.5.
 28. The apparatus as set forthin claim 25, wherein said apparatus generates shaft power from saidfirst output shaft at a simple cycle efficiency of at least about 52percent, based on the ratio of fuel energy input to mechanical energyoutput from said first output shaft, when operating at an inlet velocityof at least Mach
 4. 29. The apparatus as set forth in claim 6, or claim7, or claim 25, wherein said apparatus generates electrical power with acombined cycle efficiency of at least 65 percent based on the ratio offuel energy input to electrical energy output, when operating at aninlet velocity of at least Mach 3.5.
 30. The apparatus as set forth inclaim 1, or claim 25, wherein said apparatus consumes between about3,700 to about 4,200 BTU/Hp-hr, based on combined cycle efficiency andthe ratio of fuel energy input to electrical energy output.
 31. Anapparatus for generating power, comprising: (a) a support structuremeans, said support structure means comprising (i) an oxidant supplymeans, and (ii) a first housing means, said first housing means furthercomprising a first rotor side surface, and (iii) a second housing means,said second housing means further comprising a second rotor sidesurface; (b) an output shaft means, said output shaft means rotatablysecured with respect to said support structure means; (c) a rotor means,said rotor means (i) has a radius R, (ii) is connected to said outputshaft means to provide rotary motion of said output shaft means uponrotation of said rotor means, (ii) comprises at least one materialhaving a specific strength capability in excess of 683,220 inches, (iii)comprises a first surface portion, said first surface portion rotatablypositioned in a close fitting, first spaced apart relationship adjacentto said first rotor side surface of said first housing means, and (iv)comprises a second surface portion, said second surface portionrotatably positioned in a close fitting, second spaced apartrelationship adjacent to said second rotor side surface of said secondhousing means, and (v) wherein each of said first and said second spacedapart relationships are defined by a gap width “s” which is smallcompared to radius “R” of said rotor, to at least a partially house saidrotor means in a tight fitting relationship, so as to minimizeaerodynamic drag on said rotor means; (d) one or more ramjet thrustmeans, said one or more ramjet thrust means (i) each secured to saidrotor means for rotation therewith, (ii) each further comprise an inletand an outlet, and wherein said inlet and said outlet are substantiallyaligned in a linear configuration with respect to an inlet airflow,(iii) each further comprise a shaped external portion, said shapedexternal portion comprising a substantially constant cross-sectionalsize and substantially constant cross-sectional shape when sequentiallyexamined in cross-section perpendicular to the axis of an inlet airflowfrom a forward cross-section to a rearward cross-section, to therebyminimize pressure drag when said one or more ramjet thrust means operateat an inlet airflow velocity M₀ of at least Mach 1.5, and (iv) each ofwhich mixes fuel supplied thereto with an oxidant supplied via saidoxidant supply means, to burn said fuel to generate hot combustion gaswhich escapes from said one or more ramjet thrust means, producingthrust, thereby propelling said one or more ramjet thrust means to turnsaid rotor means and said output shaft means, thus providing shaft poweroutput from said apparatus.
 32. An apparatus for generating power,comprising: (a) a support structure, said support structure comprising(i) an oxidant supply conduit, and (ii) a first housing portion with arotor side surface, and (iii) a second housing portion with a rotor sidesurface; (b) a first output shaft, said first output shaft rotatablysecured with respect to said support structure; (c) a rotor, said rotorcomprising two or more bi-plane rotor portions, wherein said two or morebi-plane rotor portions (i) each are secured to said first output shaftto provide rotary motion of said first output shaft upon rotation ofsaid rotor bi-plane rotor portions, (ii) each further comprising (A) anupper, downwardly extending generally triangular portion, and (B) amatching lower, upwardly extending generally triangular portion, andwherein said upper portion and said lower portion of said bi-plane rotorforms therebetween an airflow inlet which receives airflow at supersonicspeed and passes at least a portion of said airflow between said upperportion and said lower portion of said bi-plane rotor, said airflowinlet contoured and substantially aligned, by shape and size of saidupper portion and said lower portion, that cancellation of an inletshock resulting from said airflow at supersonic speed is achievedbetween said upper portion and said lower portion of said bi-plane rotorportions; (d) one or more ramjet thrust modules, said one or more ramjetthrust modules (i) each secured to one of said bi-plane rotors forrotation therewith, (ii) each further comprising an inlet and an outlet,and wherein said inlet and said outlet are substantially aligned in alinear configuration with respect to an inlet airflow, (iii) eachfurther comprising a shaped external portion, said shaped externalportion comprising a substantially constant cross-sectional size whensequentially examined in cross-section perpendicular to the axis of aninlet airflow from a forward cross-section to a rearward cross-section,to thereby minimize pressure drag when said one or more ramjet thrustmodules operate at an inlet airflow velocity M₀ of at least Mach 1.5,and (iv) each of which mixes fuel supplied thereto with an oxidantsupplied via said oxidant supply conduit in said support structure, toburn said fuel to generate hot combustion gas which escapes from saidone or more ramjet thrust modules, producing thrust, thereby propellingsaid one or more ramjet thrust modules to turn said rotor and said firstoutput shaft, thus providing shaft power output from said apparatus. 33.An apparatus for generating power, comprising: (a) a support structure,said support structure comprising (i) an oxidant supply conduit, and(ii) a first housing portion with a rotor side surface, and (iii) asecond housing portion with a rotor side surface; (b) a first outputshaft, said first output shaft rotatably secured with respect to saidsupport structure; (c) a rotor, said rotor comprising (i) a centraldisc, and (ii) two or more bi-plane rotor portions radially extendingfrom said central disc, wherein said two or more bi-plane rotor portions(A) each are secured to said first output shaft to provide rotary motionof said first output shaft upon rotation of said rotor bi-plane rotorportions, (B) each further comprising (1) an upper, downwardly extendinggenerally triangular portion, and (2) a matching lower, upwardlyextending generally triangular portion, and (3) wherein said upperportion and said lower portion of said bi-plane rotor forms therebetweenan airflow inlet which receives airflow at supersonic speed and passesat least a portion of said airflow between said upper portion and saidlower portion of said bi-plane rotor, said airflow inlet contoured andsubstantially aligned, by shape and size of said upper portion and saidlower portion, that cancellation of an inlet shock resulting from saidairflow at supersonic speed is achieved between said upper portion andsaid lower portion of said bi-plane rotor portions; (d) one or moreramjet thrust modules, said one or more ramjet thrust modules (i) eachsecured to one of said bi-plane rotors for rotation therewith, (ii) eachfurther comprising an inlet and an outlet, and wherein said inlet andsaid outlet are substantially aligned in a linear configuration withrespect to an inlet airflow, (iii) each further comprising a shapedexternal portion, said shaped external portion comprising asubstantially constant cross-sectional size when sequentially examinedin cross-section perpendicular to the axis of an inlet airflow from aforward cross-section to a rearward cross-section, to thereby minimizepressure drag when said one or more ramjet thrust modules operate at aninlet airflow velocity M₀ of at least Mach 1.5, and (iv) each of whichmixes fuel supplied thereto with an oxidant supplied via said oxidantsupply conduit in said support structure, to burn said fuel to generatehot combustion gas which escapes from said one or more ramjet thrustmodules, producing thrust, thereby propelling said one or more ramjetthrust modules to turn said rotor and said first output shaft, thusproviding shaft power output from said apparatus.
 34. The apparatus asset forth in claim 32, or in claim 33, wherein said shaped externalportion of said one or more ramjet thrust modules comprises asubstantially constant cross-sectional shape, when sequentially examinedin cross-section perpendicular to the axis of an inlet airflow from aforward cross-section to a rearward cross-section, to thereby minimizepressure drag when operating at an inlet airflow velocity M₀ of at leastMach 1.5.
 35. The apparatus as set forth in claim 32, or claim 33,wherein at least one material comprising said rotor has a specificstrength in excess of 683,220 inches.
 36. The apparatus as set forth inclaim 1, or in claim 25, or in claim 32, or in claim 33, wherein atleast one material comprising said rotor has a specific strength between683,220 inches and 1,300,250 inches.
 37. The apparatus as set forth inclaim 1, or in claim 25, or in claim 32, or in claim 33, wherein atleast a portion of material comprising said rotor has a specificstrength of approximately 1,300,250 inches.
 38. The apparatus as setforth in claim 1, or claim 25, or claim 32, or 33, wherein at least aportion of material comprising said rotor has a specific strength inexcess of 1,300,250 inches.
 39. The apparatus as set forth in claim 1,or claim 25, or claim 32, or claim 33, wherein at least a portion ofmaterial comprising said rotor has a specific strength in the range fromabout 1,300,250 inches to about 3,752,600 inches.
 40. The apparatus asset forth in claim 1, or claim 25, or claim 32, or claim 33, wherein atleast a portion of material comprising said rotor has a specificstrength of about 3,752,600 inches.
 41. The apparatus as set forth inclaim 1, or claim 25, or claim 32, or claim 33, wherein at least aportion of material comprising said rotor has a specific strength inexcess of 3,752,600 inches.
 42. The apparatus as set forth in claim 1,or claim 25, or claim 32, or claim 33, wherein at least a portion ofmaterial comprising said rotor has a specific strength between 3,752,600inches and 15,000,000 inches.
 43. The apparatus as set forth in claim 1,or claim 25, or claim 32, or claim 33, wherein at least a portion ofmaterial comprising said rotor has a specific strength of about15,000,000 inches.
 44. An apparatus for generating power as set forth inclaim 1, or claim 25, or claim 32, or claim 33, wherein said at leastone ramjet thrust module operates at an inlet velocity M₀ of betweenabout Mach 1.5 and Mach 2.0.
 45. An apparatus for generating power asset forth in claim 1, or claim 25, or claim 32, or claim 33, whereinsaid at least one ramjet thrust module operates at an inlet velocity M₀of at least Mach 2.0.
 46. An apparatus for generating power as set forthin claim 1, or claim 25, or claim 32, or claim 33, wherein said at leastone ramjet thrust module operates at an inlet velocity M₀ of at leastMach 2.5.
 47. An apparatus for generating power as set forth in claim 1,or claim 25, or claim 32, or claim 33, wherein said at least one ramjetthrust module operates at an inlet velocity M₀ of at least Mach 3.0. 48.An apparatus for generating power as set forth in claim 1, or claim 25,or claim 32, or claim 33, wherein said at least one ramjet thrust moduleoperates at an inlet velocity M₀ between Mach 3.0 and Mach 4.5.
 49. Anapparatus for generating power as set forth in claim 1, or claim 25, orclaim 32, or claim 33, wherein said at least one ramjet thrust moduleoperates at an inlet velocity M₀ of approximately Mach 3.5.
 50. Anapparatus for generating power, as set forth in claim 32, or in claim33, further comprising a heat recovery section, said heat recoverysection arranged to receive said hot combustion gas from said one ormore ramjet thrust modules, said heat recovery section furthercomprising an inlet, an outlet, and a secondary working fluid forcirculation to and from said heat recovery section, whereby said hotcombustion gas is cooled by recovering heat therefrom and transferringsuch recovered heat into said secondary working fluid.
 51. The apparatusas set forth in claim 50, wherein said secondary working fluid is usedto provide thermal energy.
 52. The apparatus as set forth in claim 50,wherein said secondary working fluid comprises water, and wherein uponheating of said secondary working fluid, steam is produced.
 53. Theapparatus of claim 52, further comprising a steam turbine, said steamturbine having a steam turbine output shaft, wherein said steam whichresults from heating of said water is contained under pressure and fedto said steam turbine to produce useful work on said steam turbineoutput shaft.
 54. The apparatus as set forth in claim 53, wherein saidsteam turbine output shaft is operatively connected to a firstelectrical generator, and wherein said useful work on said steam turbineoutput shaft turns said first electrical generator to produceelectricity.
 55. The apparatus as set forth in claim 53, furthercomprising a second electrical generator, and wherein said shaft workproduced by said steam turbine output shaft turns said second electricalgenerator to produce electric power.
 56. The apparatus as set forth inclaim 32, or in claim 33, wherein said first output shaft is operativelyconnected to a first electrical generator, and wherein said mechanicalwork provided at said first output shaft turns said first electricalgenerator to produce electricity.
 57. The apparatus as set forth inclaim 32, or in claim 33, wherein said apparatus generates shaft powerat a simple cycle efficiency of at least 37 percent based on the ratioof fuel energy input to mechanical energy output, when operating at aninlet velocity of at least Mach
 3. 58. The apparatus as set forth inclaim 32, or in claim 33, wherein said apparatus generates shaft powerat a simple cycle efficiency of at least about 45 percent, based on theratio of fuel energy input to mechanical energy output, when operatingat an inlet velocity of at least Mach 3.5.
 59. The apparatus as setforth in claim 32, or in claim 33, wherein said apparatus generatesshaft power at a simple cycle efficiency of at least 52 percent, basedon the ratio of fuel energy input to mechanical energy output, whenoperating at an inlet velocity of at least Mach
 4. 60. The apparatus asset forth in claims 32 or in claim 33, wherein said apparatus generateselectrical power with a combined cycle efficiency of at least 65 percentbased on the ratio of fuel energy input to electrical energy output,when operating at an inlet velocity of at least Mach 3.5.
 61. Theapparatus as set forth in claim 32, or in claim 33, wherein saidapparatus consumes less than about 7,000 BTU/Hp-hr, based on simplecycle operation and the ratio of fuel energy input to mechanical energyoutput.
 62. The apparatus as set forth in claim 32, or in claim 33,wherein said apparatus consumes less than about 5,700 BTU/Hp-hr, basedon simple cycle operation and the ratio of fuel energy input tomechanical energy output.
 63. The apparatus as set forth in claim 32, orin claim 33, wherein said apparatus operates at in inlet velocity ofapproximately Mach 3.5, and wherein said apparatus consumes betweenabout 5,500 to about 5,700 BTU/Hp-hr, based on simple cycle operationand the ratio of fuel energy input to mechanical energy output.
 64. Theapparatus as set forth in claims 32, or in claim 33, wherein saidapparatus consumes less than about 4,200 BTU/Hp-hr, based on combinedcycle operation and the ratio of fuel energy input to electrical energyoutput.
 65. The apparatus as set forth in claims 32, or in claim 33,wherein said apparatus consumes less than about 4,000 BTU/Hp-hr, basedon combined cycle operation and the ratio of fuel energy input toelectrical energy output.
 66. The apparatus as set forth in claim 32, orin claim 33, wherein said apparatus consumes between about 3,700 toabout 4,200 BTU/Hp-hr, based on combined cycle efficiency and the ratioof fuel energy input to electrical energy output.
 67. The apparatus ofclaim 33, or of claim 23, or of claim 24, wherein said central disc istapered.
 68. The apparatus of claim 32, or of claim 33, or of claim 23,or of claim 24, wherein said bi-plane is tapered.
 69. The apparatus ofclaim 1, or of claim 25, or of claim 32, or of claim 33, wherein saidrotor comprises a metal matrix composite.
 70. The apparatus of claim 69,wherein said metal matrix composite comprises titanium.
 71. Theapparatus of claim 69, wherein said metal matrix composite comprisessilicon carbide.
 72. The apparatus of claim 69, wherein said metalmatrix composite comprises carbon fibers.
 73. The apparatus of claim 69,wherein said metal matrix composite further comprises silicon carbidefilaments.
 74. The apparatus of claim 1, or claim 25, or claim 32, orclaim 33, wherein said rotor comprises silicon carbide coated carbonfibers embedded in a titanium metal substrate.
 75. The apparatus ofclaim 1, or claim 25, or claim 32, or claim 33, wherein said rotorfurther comprises high strength fiber windings.
 76. The apparatus ofclaim 75, wherein said high strength fiber windings comprisemonofilament carbon fibers.
 77. The apparatus of claim 75, wherein saidhigh strength fiber windings comprise kevlar fibers.
 78. The apparatusof claim 75, wherein said high strength fiber windings each have abeginning and an end, and wherein said beginning and said end are eachsecured at an intermediate radial position in said rotor.
 79. Theapparatus of claim 75, wherein said high strength fiber windings are (a)spread vertically in a rotor thickness direction, and (b) are spreadhorizontally in a rotor leading edge to trailing edge direction, and (c)wherein said high strength fiber windings are provided in aconfiguration wherein said windings extend further in the leading edgeto trailing edge direction than in the rotor thickness direction. 80.The apparatus of claim 1, or claim 32, or claim 33, further comprisingan electrical generator, said generator operatively connected to saidrotor, so that rotation of said rotor energizes said generator, tothereby generate electrical power.
 81. The apparatus of claim 1, orclaim 32, or claim 33, further comprising a reaction turbine, saidreaction turbine comprising a plurality of aerodynamically shaped bladeportions adapted to react to high velocity exhaust gases impingingthereagainst, so as to react thereto and to thereby turn said reactionturbine to produce useful work.
 82. The apparatus of claim 81, whereinsaid reaction turbine is annular in shape.
 83. The apparatus of claim81, wherein said reaction turbine is operatively connected to anelectrical generator, so that work produced by said reaction turbine isutilized to produce electrical energy.
 84. The apparatus of claim 1, orof claim 32, or of claim 33, further comprising (a) a starter motor; and(b) a gear box; (c) wherein said starter motor is operatively connectedto said rotor through said gear box, and wherein said starter motor isused to provide power to said rotor to rotate said rotor and theaccompanying one or more ramjet thrust modules until said one or moreramjet thrust modules reach an inlet speed which enables said one ormore ramjet thrust modules to begin oxidation of fuel and to provide itsown thrust for sustaining rotation of said rotor.
 85. The apparatus ofclaim 84, wherein said starter motor is alternately usable as (a) astarter with electrical energy supplied thereto, and (b) a generator, sothat said starter motor is used to generate electrical power therefromas said rotor is turned by rotational energy supplied by said one ormore ramjet thrust modules.
 86. The apparatus of claim 1, or of claim32, or of claim 33, wherein said one or more ramjet thrust modulescomprises a mixed compression type inlet, so that said one or moreramjet thrust modules are self-starting.
 87. The apparatus of claim 1,or of claim 32, or of claim 33, wherein each of said one or more ramjetthrust modules further comprises an internal compression inlet.
 88. Theapparatus of claim 1, or of claim 32, or of claim 33, wherein each ofsaid one or more ramjet thrust modules further comprise exhaust gasoutlet portions having a preselected output gas reaction angle alpha(α), said angle alpha (α) being at least five degrees outwardly from thetangent to the circle of rotation of each of said one or more ramjetthrust modules, so as to direct exhaust gases outwardly from each ofsaid one or more ramjet thrust modules.
 89. An apparatus for generatingpower, comprising: (a) a support structure, said support structurecomprising (i) an oxidant supply conduit, and (ii) a first housingportion with a rotor side surface, and (iii) a second housing portionwith a rotor side surface; (b) a first output shaft, said first outputshaft rotatably secured with respect to said support structure; (c) arotor, wherein said rotor (i) is connected to said first output shaft toprovide rotary motion of said first output shaft upon rotation of saidrotor, (ii) comprises a first surface portion, said first surfaceportion rotatably positioned in a close fitting, first spaced apartrelationship adjacent to said rotor side surface of said first housingportion, and (iii) comprises a second surface portion, said secondsurface portion rotatably positioned in a close fitting, second spacedapart relationship adjacent to said rotor side surface of said secondhousing portion, and (iv) wherein each of said first and said secondspaced apart relationships are defined by a gap width “s” which is smallcompared to radius “R” of said rotor, to at least a partially house saidrotor in a tight fitting relationship, so as to minimize aerodynamicdrag on said rotor; (d) one or more ramjet thrust modules, said one ormore ramjet thrust modules (i) each secured to said rotor for rotationtherewith, (ii) each further comprising an inlet and an outlet, andwherein said inlet and said outlet are substantially aligned in a linearconfiguration with respect to an inlet airflow, (iii) each furthercomprising a shaped external portion, said shaped external portioncomprising (A) a substantially constant cross-sectional size, and (B) asubstantially constant cross-sectional shape, when sequentially examinedin cross-section perpendicular to the axis of an inlet airflow from aforward cross-section to a rearward cross-section, (C) to therebyminimize pressure drag when said one or more ramjet thrust modulesoperate at an inlet airflow velocity M₀ of at least about Mach 2.0, and(iv) each of the one or more ramjet thrust modules mixes fuel suppliedthereto with an oxidant supplied via said oxidant supply passageway insaid support structure, to burn said fuel to generate hot combustion gaswhich escapes from said one or more ramjet thrust modules, therebyproducing thrust and propelling said one or more ramjet thrust modulesand turning said rotor and said first output shaft, thus providing shaftpower output; (e) a heat recovery section, said heat recovery sectionarranged to receive said hot combustion gas from said one or more ramjetthrust modules, said heat recovery section further comprising (i) a heatrecovery inlet, (ii) a heat recovery outlet, and (iii) a secondaryworking fluid for circulation to and from said heat recovery section,(iv) whereby said hot combustion gas is cooled by recovering heattherefrom and transferring such recovered heat into said secondaryworking fluid; (f) a first electrical generator, said first electricalgenerator operatively connected to said first output shaft, and whereinsaid mechanical work provided at said first output shaft turns saidfirst electrical generator to produce electricity.